Unducted single rotor engine and method for operation

ABSTRACT

A propulsion system is provided, the propulsion system including a rotor assembly configured to rotate relative to the engine centerline axis, and wherein one or more blades of the rotor assembly are configured to rotate along a blade pitch angle axis. A vane assembly is positioned in aerodynamic relationship with the rotor assembly. The vane assembly includes one or more vanes, wherein each vane includes a vane pitch angle. A controller is configured to execute operations, the operations including moving each blade to a reverse thrust position about its respective blade pitch axis, and adjusting each vane about its respective vane pitch axis when the plurality of blades is in the reverse thrust position to modify an amount of reverse thrust generated by the propulsion system.

CROSS-REFERENCE TO RELATED APPLICATIONS

This application is a non-provisional application claiming the benefitof priority under 35 U.S.C. § 119(e) to U.S. Provisional Application No.62/915,364, filed Oct. 15, 2019, which is hereby incorporated byreference in its entirety.

FIELD

This application is generally directed to a turbomachine engine,including architectures for such an engine and methods for operating forsuch an engine.

BACKGROUND

A turbofan engine operates on the principle that a central gas turbinecore drives a bypass fan, the bypass fan being located at a radiallocation between a nacelle of the engine and the engine core. With sucha configuration, the engine is generally limited in a permissible sizeof the bypass fan, as increasing a size of the fan correspondinglyincreases a size and weight of the nacelle.

An open rotor engine, by contrast, operates on the principle of havingthe bypass fan located outside of the engine nacelle. This permits theuse of larger rotor blades able to act upon a larger volume of air thanfor a traditional turbofan engine, potentially improving propulsiveefficiency over conventional turbofan engine designs.

Engines with an open rotor design having a fan provided by twocontra-rotating rotor assemblies have been studied. Each rotor assemblycarries an array of airfoil blades located outside the engine nacelle.As used herein, “contra-rotational relationship” means that the bladesof the first and second rotor assemblies are arranged to rotate inopposing directions to each other. Typically, the blades of the firstand second rotor assemblies are arranged to rotate about a common axisin opposing directions, and are axially spaced apart along that axis.For example, the respective blades of the first rotor assembly andsecond rotor assembly may be co-axially mounted and spaced apart, withthe blades of the first rotor assembly configured to rotate clockwiseabout the axis and the blades of the second rotor assembly configured torotate counter-clockwise about the axis (or vice versa). Contra-rotatingrotor assemblies however present technical challenges in transmittingpower from the power turbine to drive the blades of the rotor assembliesrotating in opposing directions.

BRIEF DESCRIPTION

Aspects and advantages of the invention will be set forth in part in thefollowing description, or may be obvious from the description, or may belearned through practice of the invention.

An aspect of the present disclosure is directed to a propulsion systemincluding a rotor assembly configured to rotate relative to the enginecenterline axis. One or more blades of the rotor assembly are configuredto rotate along a blade pitch angle axis. A vane assembly is positionedin aerodynamic relationship with the rotor assembly. The vane assemblyincludes one or more vanes, wherein each vane includes a vane pitchangle. A controller is configured to execute operations, the operationsincluding moving each blade of the plurality of blades to a reversethrust position about its respective blade pitch axis, wherein a leadingedge of each blade is located aft of a trailing edge of the respectiveblade at a radial span location when in the reverse thrust position, andadjusting each vane of the plurality of vanes about its respective vanepitch axis when the plurality of blades is in the reverse thrustposition to modify an amount of reverse thrust generated by thepropulsion system.

These and other features, aspects and advantages of the presentinvention will become better understood with reference to the followingdescription and appended claims. The accompanying drawings, which areincorporated in and constitute a part of this specification, illustrateembodiments of the invention and, together with the description, serveto explain the principles of the invention.

BRIEF DESCRIPTION OF THE DRAWINGS

A full and enabling disclosure of the present invention, including thebest mode thereof, directed to one of ordinary skill in the art, is setforth in the specification, which makes reference to the appendedfigures, in which:

FIG. 1 is a cross-sectional side view of an embodiment of a propulsionsystem according to an aspect of the present disclosure;

FIG. 2 is a flowpath view of an exemplary embodiment of a rotor assemblyof the propulsion system of FIG. 1;

FIG. 3 is a perspective view of an exemplary embodiment of the rotorassembly of FIG. 2;

FIGS. 4-5 are top-down views of a portion of an exemplary embodiment ofa variable blade pitch rotor assembly for a propulsion system accordingto an aspect of the present disclosure;

FIGS. 6-7 are schematic depictions of modes of operation for theexemplary embodiments of the variable blade pitch rotor assembly ofFIGS. 4-5;

FIG. 8 is an exemplary embodiment of a vane of a vane assembly of thepropulsion system of FIG. 1;

FIGS. 9-13 are roll-out views of embodiments of the vane assembly of thepropulsion system of FIG. 1;

FIG. 14 is an exemplary embodiment of positions of an articulatable vaneof the propulsion system of FIG. 1;

FIGS. 15-16 are top-down views of a portion of an exemplary embodimentof a variable vane pitch assembly for a propulsion system according toaspects of the present disclosure;

FIGS. 17-21 are top-down schematic views depicting operations of anexemplary embodiment of a blade and a vane of the propulsion system ofFIG. 1;

FIG. 22 is a flowchart outlining steps of a method for adjusting thrustvector for an unducted rotor engine;

FIGS. 23-29 are schematic depictions of embodiments of computing systemsconfigured to operate one or more propulsion systems according toaspects of the present disclosure; and

FIGS. 30-31 are schematic depictions of embodiments of enginearrangements and computing systems according to embodiments depicted inFIGS. 23-29.

Repeat use of reference characters in the present specification anddrawings is intended to represent the same or analogous features orelements of the present invention.

DETAILED DESCRIPTION

Reference will now be made in detail to present embodiments of theinvention, one or more examples of which are illustrated in theaccompanying drawings. The detailed description uses numerical andletter designations to refer to features in the drawings. Like orsimilar designations in the drawings and description have been used torefer to like or similar parts of the invention.

The word “exemplary” is used herein to mean “serving as an example,instance, or illustration.” Any implementation described herein as“exemplary” is not necessarily to be construed as preferred oradvantageous over other implementations.

As used herein, the terms “first”, “second”, and “third” may be usedinterchangeably to distinguish one component from another and are notintended to signify location or importance of the individual components.

The terms “forward” and “aft” refer to relative positions within a gasturbine engine or vehicle, and refer to the normal operational attitudeof the gas turbine engine or vehicle. For example, with regard to a gasturbine engine, forward refers to a position closer to an engine inletand aft refers to a position closer to an engine nozzle or exhaust.

The terms “upstream” and “downstream” refer to the relative directionwith respect to fluid flow in a fluid pathway. For example, “upstream”refers to the direction from which the fluid flows, and “downstream”refers to the direction to which the fluid flows.

The terms “coupled,” “fixed,” “attached to,” and the like refer to bothdirect coupling, fixing, or attaching, as well as indirect coupling,fixing, or attaching through one or more intermediate components orfeatures, unless otherwise specified herein.

The singular forms “a”, “an”, and “the” include plural references unlessthe context clearly dictates otherwise.

Approximating language, as used herein throughout the specification andclaims, is applied to modify any quantitative representation that couldpermissibly vary without resulting in a change in the basic function towhich it is related. Accordingly, a value modified by a term or terms,such as “about”, “approximately”, and “substantially”, are not to belimited to the precise value specified. In at least some instances, theapproximating language may correspond to the precision of an instrumentfor measuring the value, or the precision of the methods or machines forconstructing or manufacturing the components and/or systems. Forexample, the approximating language may refer to being within a 1, 2, 4,10, 15, or 20 percent margin, unless otherwise specified.

Here and throughout the specification and claims, range limitations arecombined and interchanged, such ranges are identified and include allthe sub-ranges contained therein unless context or language indicatesotherwise. For example, all ranges disclosed herein are inclusive of theendpoints, and the endpoints are independently combinable with eachother.

References to “noise”, “noise level”, or “perceived noise”, orvariations thereof, are understood to include perceived noise levels,effective perceived noise levels EPNL, instantaneous perceived noiselevels PNL(k), or tone-corrected perceived noise levels PNLT(k), or oneor more duration correction factors, tone correction factors, or otherapplicable factors, as defined by the Federal Aviation Administration(FAA), the European Union Aviation Safety Agency (EASA), theInternational Civil Aviation Organization (ICAO), Swiss Federal Officeof Civil Aviation (FOCA), or committees thereof, or other equivalentregulatory or governing bodies. Where certain ranges of noise levels(e.g., in decibels, or dB) are provided herein, it will be appreciatedthat one skilled in the art will understand methods for measuring andascertaining of such levels without ambiguity or undue experimentation.Methods for measuring and ascertaining one or more noise levels asprovided herein by one skilled in the art, with reasonable certainty andwithout undue experimentation, include, but are not limited to,understanding of measurement systems, frames of reference (including,but not limited to, distances, positions, angles, etc.) between theengine and/or aircraft relative to the measurement system or otherperceiving body, or atmospheric conditions (including, but not limitedto, temperature, humidity, dew point, wind velocity and vector, andpoints of reference for measurement thereof), as may be defined by theFAA, EASA, ICAO, FOCA, or other regulatory or governing body.

As used herein, the terms “processor” and “computer,” and related terms,e.g., “processing device,” “computing device,” and “controller”, are notlimited to just those integrated circuits referred to in the art as acomputer, but further broadly refers to one or more processing devicesincluding one or more of a microcontroller, a microcomputer, aprogrammable logic controller (PLC), an application specific integratedcircuit, and other programmable circuits, and these terms are usedinterchangeably herein. In the embodiments described herein, thecomputer or controller may additionally include memory. The memory mayinclude, but is not limited to, a computer-readable medium, such as arandom access memory (RAM), a computer-readable non-volatile medium,such as a flash memory. Alternatively, a floppy disk, a compactdisc-read only memory (CD-ROM), a magneto-optical disk (MOD), and/or adigital versatile disc (DVD) may also be used. Also, in the embodimentsdescribed herein, the computer or controller may include one or moreinput channels and/or one or more output channels. The input channelsmay be, but are not limited to, computer peripherals associated with anoperator interface such as a mouse and a keyboard, or sensors, such asengine sensors associated with an engine, such as a gas turbine engine,for determining operating parameters of the engine. Furthermore, in theexemplary embodiment, the output channels may include, but are not belimited to, an operator interface monitor, or the output channels may belinked to various components to control such components based, e.g., ondata reviewed from the input channels and/or data or instructions storedin the memory. For example, the memory may store software or otherinstructions, which when executed by the controller or processor allowthe controller to perform certain operations or functions, such as oneor more of the operations or functions for which the controller isconfigured and/or one or more of the methods described herein. The term“software” may include any computer program stored in memory, oraccessible by the memory, for execution by, e.g., the controller,processor, clients, and servers.

Referring to FIG. 1, in general, embodiments of an engine 10 variouslydepicted and described herein include a computing system 210 configuredto include one or more controllers 1600, 1610, 1700, 1800 depicted anddescribed herein, and/or configured to execute steps of a method orother operations provided herein. The computing system 210 cancorrespond to any suitable processor-based device, including one or morecomputing devices, such as described above. For instance, FIG. 1illustrates one embodiment of suitable components that can be includedwithin the computing system 210. As shown in FIG. 1, the computingsystem 210 can include a processor 212 and associated memory 214configured to perform a variety of computer-implemented functions (e.g.,performing the methods, steps, calculations and the like disclosedherein).

As shown, the computing system 210 can include control logic 216 storedin memory 214. The control logic 216 may include instructions that whenexecuted by the one or more processors 212 cause the one or moreprocessors 212 to perform operations, such as the method or operationsdescribed above, such as in regard to controllers 1600, 1610, 1700, 1800depicted and described herein. Additionally, as shown in FIG. 1, thecomputing system 210 can also include a communications interface module230. In several embodiments, the communications interface module 230 caninclude associated electronic circuitry that is used to send and receivedata. As such, the communications interface module 230 of the computingsystem 210 can be used to send and/or receive data to/from engine 10 andthe compressor section 21. In addition, the communications interfacemodule 230 can also be used to communicate with any other suitablecomponents of the engine 10, including any number of motors, actuators,linkages, vane or blade pitch change mechanisms, sensors, or otheractuatable structures, such as one or more of those depicted anddescribed herein.

It should be appreciated that the communications interface module 230can be any combination of suitable wired and/or wireless communicationsinterfaces and, thus, can be communicatively coupled to one or morecomponents of the compressor section 21 or the engine 10 via a wiredand/or wireless connection. As such, the computing system 210 mayobtain, determine, store, generate, transmit, or operate any one or moresteps of the operations such as described herein with regard to theengine 10 or an apparatus (e.g., aircraft or other vehicle) to which theengine 10 is attached.

Further, certain embodiments of the unducted single rotor turbomachineengine described hereinbelow may include an electric machine. Anelectric machine may generally include a stator and a rotor, the rotorrotatable relative to the stator. Additionally, the electric machine maybe configured in any suitable manner for converting mechanical power toelectrical power, or electrical power to mechanical power. For example,the electric machine may be configured as an asynchronous or inductionelectric machine operable to generate or utilize alternating current(AC) electric power. Alternatively, the electric machine may beconfigured as a synchronous electric machine operable to generate orutilize AC electric power or direct current (DC) electric power. In sucha manner it will be appreciated that the stator, the rotor, or both maygenerally include one or more of a plurality of coils or windingarranged in any suitable number of phases, one or more permanentmagnets, one or more electromagnets, etc.

It would be desirable to provide an open rotor propulsion systemutilizing a single rotating rotor assembly analogous to a traditionalturbofan engine bypass fan which reduces the complexity of the design,yet yields a level of propulsive efficiency comparable tocontra-rotating propulsion designs with a significant weight and lengthreduction.

Embodiments of a single unducted rotor engine 10 are provided herein.Embodiments of the engine, propulsion system, or thrust-producing systemprovided herein may generate an increased unducted rotor efficiency at,and above a threshold power loading (i.e., power/area of rotor airfoil).In certain embodiments, the threshold power loading is 25 horsepower perft² or greater at cruise altitude. In particular embodiments of theengine, structures and methods provided herein generate power loadingbetween 25 horsepower/ft² and 100 horsepower/ft² at cruise altitude.Cruise altitude is generally an altitude at which an aircraft levelsafter climb and prior to descending to an approach flight phase. Invarious embodiments, the engine is applied to a vehicle with a cruisealtitude up to approximately 65,000 ft. In certain embodiments, cruisealtitude is between approximately 28,000 ft and approximately 45,000 ft.In still certain embodiments, cruise altitude is expressed in flightlevels based on a standard air pressure at sea level, in which a cruiseflight condition is between FL280 and FL650. In another embodiment,cruise flight condition is between FL280 and FL450. In still certainembodiments, cruise altitude is defined based at least on a barometricpressure, in which cruise altitude is between approximately 4.85 psiaand approximately 0.82 psia based on a sea level pressure ofapproximately 14.70 psia and sea level temperature at approximately 59degree Fahrenheit. In another embodiment, cruise altitude is betweenapproximately 4.85 psia and approximately 2.14 psia. It should beappreciated that in certain embodiments, the ranges of cruise altitudedefined by pressure may be adjusted based on a different reference sealevel pressure and/or sea level temperature.

Various embodiments of the single unducted rotor engine include a vaneassembly 30 in aerodynamic relationship with a bladed rotor assembly 20.Referring to FIG. 1, the vane assembly 30 is positioned aft (i.e.,proximate to aft end 99) or generally downstream (relative to normalforward operation, schematically depicted by arrow FW) of a singleunducted rotor assembly 20. The vane assembly 30 may generally define ade-swirler device configured to reduce or convert kinetic energy lossesfrom unducted rotors into thrust output. In certain embodiments, thevane assembly 30 is configured to adjust vane pitch angle 90 based atleast on output velocity vectors from the rotor assembly 20. Theadjustable vane pitch angle is configured to output a desired thrustvector based on a desired engine operation (e.g., forward thrust,neutral or no thrust, or reverse thrust) and desired acoustic noiselevel. In still certain embodiments, the bladed rotor assembly 20 isconfigured to adjust blade pitch angle 91 based at least on a desiredoutput velocity vector to the vane assembly 30, a desired engineoperation, or a desired acoustic noise level. In still variousembodiments, the rotor assembly 20 is configured to adjust rotor planebased on an angle of attack of incoming air to the rotor assembly, suchas to adjust an output velocity vector to the vane assembly and reduceor eliminate undesired noise levels from the rotor assembly.

Certain embodiments of the single unducted rotor engine 10 provide noisereduction or attenuation based on dynamic blade pitch angle changes,vane pitch angle changes, and/or rotor plane angle changes relative toangle of attack of incoming air and output air velocity from the rotorassembly to an aft vane assembly. Additionally, or alternatively,embodiments of the engine 10 provided herein may attenuate low frequencynoise, such as those that may propagate to the ground while an engine isat cruise altitude, or as may be referred to as “en-route noise.”Various embodiments of the engine are configured to desirably alterrotor plane angle, blade pitch angle, and/or vane pitch angle tomitigate propagation of undesired noise to the ground and the fuselage.Additionally, the engine 10 may be configured to desirably deflect noiseupward (e.g., skyward) rather than toward the ground. As such, perceivednoise levels may be reduced or mitigated by one or more structuresprovided herein.

Referring now to the drawings, FIG. 1 shows an elevationalcross-sectional view of an exemplary embodiment of a single unductedrotor engine 10. As is seen from FIG. 1, the engine 10 takes the form ofan open rotor propulsion system and has a rotor assembly 20 whichincludes an array of airfoil blades 21 around a longitudinal axis 11 ofengine 10. Blades 21 are arranged in typically equally spaced relationaround the longitudinal axis 11, and each blade 21 has a root 223 and atip 246 and a span defined therebetween.

Additionally, engine 10 includes a gas turbine engine having a core (orhigh speed system) 40 and a low speed system. The core engine 40generally includes a high speed compressor 4042, a high speed turbine4044, and a high speed shaft 4045 extending therebetween and connectingthe high speed compressor 4042 and high speed turbine 4044. The highspeed compressor 4042, the high speed turbine 4044, and the high speedshaft 4045 may collectively define and be referred to as a high speedspool 4046 of the engine. Further, a combustion section 4048 is locatedbetween the high speed compressor 4042 and high speed turbine 4044. Thecombustion section 4048 may include one or more configurations forreceiving a mixture of fuel and air and providing a flow of combustiongasses through the high speed turbine for driving the high speed spool4046.

The low speed system 50 similarly includes a low speed turbine 5050, alow speed compressor or booster, 5052, and a low speed shaft 5055extending between and connecting the low speed compressor 5052 and lowspeed turbine 5050. The low speed compressor 5052, the low speed turbine5050, and the low speed shaft 5055 may collectively define and bereferred to as a low speed spool 5054 of the engine.

In various embodiments, the core engine 40 may include a third-streamflowpath 1063, such as to bypass flow from a core flowpath downstream ofone or more compressors. The third-stream flowpath 1063 may generallydefine a concentric or non-concentric flowpath relative to the flowpath1062 downstream of one or more compressors or fan stages. Thethird-stream flowpath 1063 is configured to selectively remove a portionof flow from the core flowpath 1062, such as via one or more variableguide vanes, nozzles, or other actuatable flow control structures. Thethird-stream flowpath 1063 may bypass the combustion section 4048. Incertain embodiment, the third-stream flowpath 1063 furthermore bypassesall or part of the flowpath at the turbine section.

It should be appreciated that the terms “low” and “high”, or theirrespective comparative degrees (e.g., -er, where applicable), when usedwith compressor, turbine, shaft, or spool components, each refer torelative speeds within an engine unless otherwise specified. Forexample, a “low turbine” or “low speed turbine” defines a componentconfigured to operate at a rotational speed, such as a maximum allowablerotational speed, lower than a “high turbine” or “high speed turbine” atthe engine. Alternatively, unless otherwise specified, theaforementioned terms may be understood in their superlative degree. Forexample, a “low turbine” or “low speed turbine” may refer to the lowestmaximum rotational speed turbine within a turbine section, a “lowcompressor” or “low speed compressor” may refer to the lowest maximumrotational speed turbine within a compressor section, a “high turbine”or “high speed turbine” may refer to the highest maximum rotationalspeed turbine within the turbine section, and a “high compressor” or“high speed compressor” may refer to the highest maximum rotationalspeed compressor within the compressor section. Similarly, the low speedspool refers to a lower maximum rotational speed than the high speedspool. It should further be appreciated that the terms “low” or “high”in such aforementioned regards may additionally, or alternatively, beunderstood as relative to minimum allowable speeds, or minimum ormaximum allowable speeds relative to normal, desired, steady state, etc.operation of the engine.

Although the engine 10 is depicted with the low speed compressor 5052positioned forward (i.e., proximate to a forward end 98) of the highspeed compressor 4042, in certain embodiments the compressors 4042, 5052may be in interdigitated arrangement, i.e., rotary airfoils of the lowspeed compressor 5052 are in alternating arrangement along the gasflowpath with rotary airfoils of the high speed compressor 4042.Additionally, or alternatively, although the engine 10 is depicted withthe high speed turbine 4044 positioned forward of the low speed turbine5050, in certain embodiments the turbines 4044, 5050 may be ininterdigitated arrangement. Although certain embodiments or descriptionsof rotary elements provided herein may include “low pressure” or “highpressure”, it should be appreciated that the rotary elements mayadditionally, or alternatively, refer to “low speed” or “high speed”,respectively, such as based on interdigitated arrangements or otherwiseprovided above.

Referring to FIG. 1, the core engine 40 is generally encased in a cowl1056 defining a maximum diameter DM. The vane assembly 30 is extendedfrom the cowl 1056 and positioned aft of the rotor assembly 20. Invarious embodiments, the maximum diameter is defined as a flowpathsurface facing outward along the radial direction R in fluidcommunication with the flow of fluid egressed from the rotor assembly20. In certain embodiments, the maximum diameter of the cowl 1056corresponds substantially to a location or positioning of a root 335 ofa vane 31 of the vane assembly 30 extended from the cowl 1056. The rotorassembly 20 further includes a hub 1052 extended forward of theplurality of blades 21.

In certain embodiments, the engine 10 defines a length L from a forwardend 1042 of the hub 1052 to an aft end 1043 of the cowl 1056. However,it should be appreciated that the length L may correspond to an aft endof a rearward facing or pusher-configuration hub 1052 and rotor assembly20. In still certain embodiments, the engine 10 defines the length Lfrom an aft end 1043 of the cowl 1056, in which the aft end 1043 ispositioned at an egress end or exhaust 1060 of the core engine 40. Invarious embodiments, the length L may exclude a dimension of an exhaustnozzle or cap positioned radially inward of a turbomachinery flowpath1062. In various embodiments, the engine 10 includes a ratio of length(L) to maximum diameter (D_(M)) that provides for reduced installeddrag. In one embodiment, L/D_(M) is at least 2. In another embodiment,L/D_(M) is at least 2.5. In various embodiments, it should beappreciated that the L/D_(M) is for a single unducted rotor engine.

The reduced installed drag may further provide for improved efficiency,such as improved specific fuel consumption. Additionally, oralternatively, the reduced drag may provide for cruise altitude engineand aircraft operation at or above Mach 0.5. In certain embodiments, thereduced drag provides for cruise altitude engine and aircraft operationat or above Mach 0.75. In certain embodiments, such as certainembodiments of L/D_(M), the rotor assembly 20, and/or the vane assembly30 positioned aft of the rotor assembly 20, the engine 10 defines amaximum cruise altitude operating speed between approximately Mach 0.55and approximately Mach 0.85. In still particular embodiments, certainembodiments of L/D_(M), the quantity of blades at the rotor assembly 20to the quantity of vanes at the vane assembly 30, and/or the vaneassembly 30 positioned aft of the rotor assembly 20 provides the engine10 with a maximum cruise altitude operating speed between approximatelyMach 0.75 and Mach 0.85.

Moreover, it will be appreciated that the engine 10 further includes acowl 1056 surrounding the turbomachinery and defining at least in partan inlet 1058, an exhaust 1060, and the turbomachinery flowpath 1062extending between the inlet 1058 and the exhaust 1060. The inlet 1058 isfor the embodiment shown an annular or axisymmetric 360 degree inlet1058 located between the rotor assembly 20 and the vane assembly 30, andprovides a path for incoming atmospheric air to enter the turbomachineryflowpath 1062 (and compressors, combustion section, and turbines)radially inwardly of the vane assembly 30. Such a location may beadvantageous for a variety of reasons, including management of icingperformance as well as protecting the inlet 1058 from various objectsand materials as may be encountered in operation.

As is depicted, the rotor assembly 20 is driven by the turbomachinery,and more specifically, is driven by the low speed spool 5054. Morespecifically, still, engine 10 in the embodiment shown in FIG. 1includes a power gearbox 1064, and the rotor assembly 20 is driven bythe low speed spool 5054 of the turbomachinery across the power gearbox64. In such a manner, the rotating blades 21 of the rotor assembly 20may rotate around the axis 11 and generate thrust to propel the engine10, and hence an aircraft to which it is associated, in a forwarddirection FW.

The power gearbox 1064 may include a gearset for decreasing a rotationalspeed of the low speed spool 5054 relative to the low speed turbine5050, such that the rotor assembly 20 may rotate at a slower rotationalspeed than the low speed spool 5054. In certain embodiments, the powergearbox 1064 includes a gear ratio of at least 4:1. Although in variousembodiments the 4:1 gear ratio may generally provide for the low speedturbine 5050 to rotate at approximately four times the rotational speedof the rotor assembly 20, it should be appreciated that other structuresprovided herein, such as the blade pitch change mechanism and/or anelectric machine, may allow the unducted rotor assembly 20 to operatesubstantially de-coupled from the low speed turbine 5050 rotationalspeed. Moreover, when using an interdigitated counter-rotating orvaneless turbine the gear ratio may be reduced without an appreciableloss in output power from the rotor assembly 20.

Single unducted rotor engine 10 also includes in the exemplaryembodiment a vane assembly 30 which includes an array of vanes 31 alsodisposed around longitudinal axis 11, and each vane 31 has a root 335and a tip 334 and a span defined therebetween. These vanes 31 aremounted to a stationary frame and do not rotate relative to thelongitudinal axis 11. In certain embodiments, the vanes 31 include amechanism for adjusting their orientation relative to their axis 90and/or relative to the blades 21, such as further described herein. Forreference purposes, FIG. 1 also depicts a forward direction denoted witharrow FW, which in turn defines the forward and aft portions of thesystem. As shown in FIG. 1, the rotor assembly 20 is located forward ofthe turbomachinery in a “puller” configuration, and the exhaust 1060 islocated aft of the vane assembly 30.

It may be desirable that the blades 21, the vanes 31, or both,incorporate a pitch change mechanism such that the airfoils (e.g.,blades 21, vanes 31, etc.) can be rotated with respect to an axis ofpitch rotation either independently or in conjunction with one another.Such pitch change can be utilized to vary thrust and/or swirl effectsunder various operating conditions, including to provide adjust amagnitude or direction of thrust produced at the vanes 31, or to providea thrust reversing feature which may be useful in certain operatingconditions such as upon landing an aircraft, or to desirably adjustacoustic noise produced at least in part by the blades 21, the vanes 31,or aerodynamic interactions from the blades 21 relative to the vanes 31.

Vanes 31 are sized, shaped, and configured to impart a counteractingswirl to the fluid so that in a downstream direction aft of both rows ofairfoils (e.g., blades 21, vanes 31) the fluid has a greatly reduceddegree of swirl, which translates to an increased level of inducedefficiency.

Vanes 31 may have a shorter span than blades 21, as shown in FIG. 1, forexample, 50% of the span of blades 21, or may have longer span or thesame span as blades 21 as desired. Vanes 31 may be attached to anaircraft structure associated with the propulsion system, as shown inFIG. 1, or another aircraft structure such as a wing, pylon, orfuselage. Vanes 31 of the stationary element may be fewer or greater innumber than, or the same in number as, the number of blades 21 of therotating element and typically greater than two, or greater than four,in number.

In certain embodiments, the plurality of blades 21 each have a loadingdistribution such that at any location between the blade root 223 and30% blade span 246 the value of ΔRCu in the air stream is greater thanor equal to 60% of the peak ΔRCu in the air stream. Cu is thecircumferential averaged tangential velocity in a stationary frame ofreference. Vector diagrams are shown in a coordinate system in which theaxial direction is in the downward direction and tangential direction isleft to right. Multiplying the Cu times the airstream radius R gives theproperty RCu. The blade or vane loading at a given radius R is nowdefined as the change in RCu across the blade row (at a constant radiusor along a streamtube), here forth referred to as ΔRCu and is a measureof the elemental specific torque of said blade row. Desirably, the ΔRCufor the rotating element should be in the direction of rotationthroughout the span.

In certain embodiments, the blade 21 defines a more uniform ΔRCu overthe span, particularly in the region between the blade root 223 andmidspan. In fact, at a location of 30% span the value of ΔRCu is greaterthan or equal to 60% of the maximum value of ΔRCu, and, in anembodiment, is greater than or equal to 70% of the maximum value ofΔRCu, and, in an embodiment, is greater than or equal to 80% of themaximum value of ΔRCu. ΔRCu is measured across the rotor assembly 20 ina conventional manner.

In certain embodiments, a change in the blade 21 cambers in the innerportion of the blade, i.e., from about 0 to approximately 50% span, andit is expected that characteristics of exemplary embodiments could alsobe loosely defined by a camber distribution. At least one of thefollowing criteria are met: at 30% span the blade camber is at least 90%of the max camber level between 50% and 100% span; and the 0% spancamber is at least 110% of the max camber between 50% and 100% span.Embodiments of the blade 21 may include geometries or features providingloading distribution such as provided in U.S. patent Ser. No. 10/202,865B2 “Unducted Thrust Producing System” in Appendix A, and hereinincorporated by reference in its entirety for all purposes.

Blades 21 may include a metal leading edge (MLE) wrap for withstandingforeign object debris (FOD), such as bird strikes, during engineoperation. In particular embodiments, the blades 21 include a sheetmetal sheath at the leading edge. In various embodiments, the blades 21include one or more features, including orifices, voids, openings,cavities, or other frangible features configured to desirably liberateportions of the blade 21, such as to minimize damage to the fuselage ofan aircraft.

In various embodiments, the plurality of vanes 31 and/or aircraftsurfaces 1160 may include leading edge treatments such as to reduceacoustic interactions between the rotor assembly 20 and the vanes oraircraft surfaces positioned downstream of the rotor assembly 20. Thevanes 31 and/or aircraft surface 1160 may include a surface modificationelement defining a modified contour configured to decorrelate a phasedistribution of a plurality of sound sources within a source fieldpositioned on at least a portion of the vane or aircraft surface.Embodiments of the vane 31 and/or aircraft surface 1160 may includegeometries of features such as one or more surface modification elementssuch as provided in US Patent Application No. US 2017/0225773 A1 “WingLeading Edge Features to Attenuate Propeller Wake-Wing AcousticInteractions”, and herein incorporated by reference in its entirety forall purposes.

In certain embodiments, the engine 10 includes one or more of a desiredratio of blades 21 to vanes 31, a difference in a quantity of blades 21to a quantity of vanes 31, or sum of the quantity of blades 21 and thequantity of vanes 31, providing particular and unexpected benefits suchas further described herein. Furthermore, it should be appreciated thatit may be desirable to produce thrust from the rotor assembly 20depicted and described herein within one or more particular ranges ofquantity of blades 21, or more particularly, ranges of ratios,differences, and/or sums of blades 21 to vanes 31, such as to reduceinteraction noise between an unducted rotor assembly and a vaneassembly. Still further, it should be appreciated that although certainembodiments of turbo machines may provide partially overlapped ranges ofquantities of thrust-producing blades, the present disclosure providesranges, differences, or sums that, at least in part, provide a desiredthrust for an unducted rotor assembly while attenuating or mitigatingnoise produced by an interaction of the blades 21 and the vanes 31, orattenuating noise perceived by an observer or measurement device.Additionally, or alternatively, one or more vanes 31 or vane structuresdepicted and described herein may include or be configured at, at leastin part, one or more aircraft surfaces 1160 such as described herein,including, but not limited to, a wing, pylon, fuselage, empennage, ornon-wing surface.

In various embodiments, the engine 10 includes a ratio of a quantity ofblades 21 to a quantity of vanes 31 between 2:5 and 2:1, or between 2:4and 3:2, or between 0.5 and 1.5. In certain embodiments, a differencebetween the quantity of blades 21 and the quantity of vanes 31 isbetween two (2) and negative two (−2), or between one (1) and negativeone (−1). In various embodiments, the quantity of blades 21 is twenty(20) or fewer. In still certain embodiments, a sum of the quantity ofblades 21 and the quantity of vanes 31 is between twenty (20) and thirty(30), or between twenty-four (24) and twenty-eight (28), or betweentwenty-five (25) and twenty-seven (27). In one embodiment, the engine 10includes a quantity of blades 21 between eleven (11) and sixteen (16).In another embodiment, the engine 10 includes twelve (12) blades 21 andten (10) vanes 31. In still another embodiment, the engine 10 includesbetween three (3) and twenty (20) blades 21 and between three (3) andtwenty (20) vanes 31. In yet another embodiment, the engine 10 includesan equal quantity of blades 21 and vanes 31. In still yet anotherembodiment, the engine 10 includes an equal quantity of blades 21 andvanes 31, in which the quantity of blades 21 is equal to or fewer thantwenty (20). In various embodiments, the engine 10 includes acombination of the quantity of blades 21 to the quantity of vanes 31between 2:5 and 2:1, the difference between the quantity of blades 21and the quantity of vanes 31 between two (2) and negative two (−2), andthe quantity of blades 21 between eleven (11) and sixteen (16). Forexample, a difference between the quantity of blades and the quantity ofvanes may correspond to an engine having fourteen (14) blades andsixteen (16) vanes, or fourteen (14) blades and twelve (12) vanes, orsixteen (16) blades and eighteen (18) vanes, or sixteen (16) blades andfourteen (14) vanes, or eleven (11) blades and thirteen (13) vanes, oreleven (11) blades and nine (9) vanes, etc.

In particular embodiments, a combination of the vane assembly 30positioned aerodynamically aft of the blade assembly 20 to recover swirlin the flow such as described herein and the differences between thequantities of blades 21 and vanes 31 allow for decreased noise. The vaneassembly 20 being stationary relative to the engine centerline axisallows for reduced radiation efficiency of noise and redirects the noisein a manner favorable to use the difference between the quantities ofblades 21 and vanes 31 such as described herein. In contrast, enginesincluding counter-rotating unducted fan or propeller rotors withapproximately equal blade counts for the forward and aft blade rows maygenerally result in increased noise radiation compared to acounter-rotating unducted fan or propeller rotor engine including agreater difference in blade counts between the forward and aft bladerows.

It should be appreciated that embodiments of the engine 10 including oneor more ranges of ratios, differences, or sums of blades 21 to vanes 31depicted and described herein may provide advantageous improvements overturbofan or turboprop gas turbine engine configurations. In oneinstance, embodiments of the engine 10 provided herein allow for thrustranges similar to or greater than turbofan engines with largerquantities of blades or vanes, while further obviating structures suchas fan cases or nacelles. In another instance, embodiments of the engine10 provided herein allow for thrust ranges similar to or greater thanturboprop engines with similar quantities of blades, while furtherproviding reduced noise or acoustic levels such as provided herein. Instill another instance, embodiments of the engine 10 provided hereinallow for thrust ranges and attenuated acoustic levels such as providedherein while reducing weight, complexity, or issues associated with fancases, nacelles, variable nozzles, or thrust-reverser assemblies at thenacelle.

It should further be appreciated that ranges of ratios, differences,sums, and/or discrete quantities of blades 21 to vanes 31 providedherein may provide particular improvements to gas turbine engines inregard to thrust output and acoustic levels. For instance, quantities ofblades greater than those of one or more ranges provided herein mayproduce noise levels that may disable use of an open rotor engine incertain applications (e.g., commercial aircraft, regulated noiseenvironments, etc.). In another instance, quantities of blades less thanthose ranges provided herein may produce insufficient thrust output,such as to render an open rotor engine non-operable in certain aircraftapplications. In yet another instance, quantities of vanes less thanthose of one or more ranges provided herein may fail to sufficientlyproduce thrust and abate noise, such as to disable use of an open rotorengine in certain applications. In still another instance, quantities ofvanes greater than those of ranges provided herein may result inincreased weight that adversely affects thrust output and noiseabatement.

It should be appreciated that various embodiments of the single unductedrotor engine 10 depicted and described herein may allow for normalsubsonic aircraft cruise altitude operation at or above Mach 0.5. Incertain embodiments, the engine 10 allows for normal aircraft operationbetween Mach 0.55 and Mach 0.85 at cruise altitude. In still particularembodiments, the engine 10 allows for normal aircraft operation betweenMach 0.75 and Mach 0.85. In certain embodiments, the engine 10 allowsfor rotor blade tip speeds at or less than 750 feet per second (fps).

Referring now to FIGS. 2-7, in certain embodiments, the rotor assembly20 includes a variable pitch rotor blade assembly 20 having a pluralityof rotor blades 21 coupled to a disk 42 in a spaced apart manner. Asdepicted, the blades 21 extend outwardly from disk 42 generally along aradial direction R. Each of the plurality of blades 21 defines a leadingedge 244 and a tip 246 defined at a radially outer edge of eachrespective rotor blade 21. Each rotor blade 21 is also rotatablerelative to the disk 42 about a pitch axis P by virtue of the blades 21being operatively coupled to a suitable actuation assembly 48 configuredto vary the pitch of the blades 21 in a manner described in detailbelow. The blades 21, disk 42, and actuation assembly 48 are togetherrotatable about a rotor assembly longitudinal axis 12. It should beappreciated that longitudinal axis 12 may be co-axial or common to thecentral longitudinal axis 11 of the engine 10 depicted in FIG. 1.However, in other embodiments, the rotor assembly longitudinal axis 12may be offset from the engine longitudinal axis 11, such that the axes11, 12 are at an acute angle relative to one another. Additionally, incertain embodiments, the disk 42 of the variable pitch rotor assembly 20is covered by rotatable front hub 1052 aerodynamically contoured topromote an airflow through the plurality of blades 21.

Referring now to FIG. 2 the rotor assembly 20 will be described ingreater detail. FIG. 2 provides a forward-facing-aft elevational view ofthe rotor assembly 20 of the exemplary engine 10 of FIG. 1. For theexemplary embodiment depicted, the rotor assembly 20 includes twelve(12) blades 21. From a loading standpoint, such a blade count allows thespan of each blade 21 to be reduced such that the overall diameter ofrotor assembly 20 is also able to be reduced (e.g., to about twelve feetin the exemplary embodiment). That said, in other embodiments, rotorassembly 20 may have any suitable blade count and any suitable diameter,such as described herein.

Each blade 21 may have a suitable aerodynamic profile including agenerally concave pressure side and a circumferentially opposite,generally convex suction side 100. Each blade 21 extends from an innerroot end 223, which is rotatably coupled to disk 42, to a radially outerdistal tip 246. As shown, each blade 21 defines a chord length C thatextends between opposite leading edge 244 and trailing edge 245, withthe chord varying in length over the span of the blade 21.

The rotor assembly 20 also has a corresponding solidity which is aconventional parameter equal to the ratio of the blade chord BC, asrepresented by its length, divided by the circumferential pitch CP orspacing from blade to blade at the corresponding span position orradius. The circumferential pitch is equal to the circumferential lengthat the specific radial span divided by the total number of rotor bladesin the blade row. Accordingly, the solidity is directly proportional tothe number of blades and chord length and inversely proportional to thediameter.

Typical high solidity turbofan engines have adjacent blades 21 thatsubstantially overlap each other circumferentially due to the highsolidity and high stagger of the airfoils. For example, as shown in FIG.2, the blades 21 have high solidity and adjacent blades would contacteach other when passing through the flat pitch position. Due to thesolidity of the blades 21, it can be seen that the blades 21 wouldoverlap at least in region 110 if they pass through flat pitch at thesame time. In some embodiments, in order to achieve reverse thrust fromthe rotor assembly 20, it is necessary that the blades 21 pass throughflat pitch. However, given the configuration shown in FIG. 2,unacceptable blade contact will occur if the blades 21 rotate in unisonthrough flat pitch. Therefore, a rotor assembly 20 configured forasynchronous blade pitching is described below with respect to FIGS.2-7. Such a system can ensure that the blades 21 do not pass throughflat pitch at the same time, as well as provide otherperformance-related improvements to rotor assembly 20 operation, asdiscussed below.

Referring now generally to FIGS. 2-7, a pitch actuation assembly 48 inaccordance with an exemplary embodiment of the present disclosure isdepicted. As mentioned above, each blade 21 is rotatable relative to thedisk 42 about a pitch axis P. The blades 21, disk 42, and actuationassembly 48 are together rotatable about the longitudinal axis 12. Incertain embodiments, the pitch actuation assembly described in regard tothe blade pitch actuation assembly 48 is further included at one or moreof the vanes 31 of the vane assembly 30 as a vane pitch actuationassembly 148. As such, at least a portion of the pitch actuationassembly shown and described in regard to FIGS. 2-7 may be applied toone or more of the vanes 31 such as to collectively or independentlyadjust the orientation of the vane 31 about the axis 90 of eachrespective vane 31. Such independent or collective adjustment of pitchangle of the vane 31 about axis 90 may be utilized according to one ormore methods further described herein, such as one or more methods forattenuating undesired acoustic noise, for producing a desired thrustvector, and/or for producing a desired thrust load.

The actuation assembly 48 generally includes a scheduling ring 120,plurality of linkage arms 122, and one or more motors 124 (e.g., anelectric motor, a pneumatic or hydraulic actuation device, etc.). Eachblade 21 may be rotatably coupled to the disk 42 through a first end 126of a corresponding linkage arm 122 such that the first end 126 and thecorresponding blade 21 may rotate about pitch axis P (e.g., blade pitchaxis 91 in FIG. 1) relative to disk 42. In this regard, the blade 21 maybe fixedly connected to the first end 126 of the corresponding linkagearm 122, such that rotation of the linkage arm 122 causes the blade 21to rotate relative to the disk 42.

A second end 128 of the linkage arm 122 may be slidably connected to oneof the plurality of slots 130 defined in scheduling ring 120. Forexample, the second end 128 may be rotatably connected to a slidingmember 132. The sliding member 132 may be slidably received in a slot130 of the scheduling ring 120. The scheduling ring 120 is rotatableabout longitudinal centerline 12 relative to the disk 42 and isoperatively coupled with the motor 124, which is fixed relative to thedisk 42.

Each of the plurality of slots 130 on the scheduling ring 120 defines anairfoil pitch schedule. In this regard, for a given angle of rotation ofthe scheduling ring 120, the airfoil pitch schedule determines theactual pitch angle of the blades 21. In operation, the motor 124 rotatesthe scheduling ring 120 relative to the disk 42. As the scheduling ring120 rotates, sliding member 132 moves along slot 130 and the angularposition of the linkage arm 122 changes. As each linkage arm 122rotates, the corresponding blade 21 rotates as well, thus rotating eachblade 21 about pitch axis P.

Therefore, by rotating the scheduling ring 120 relative to the disk 42,each of the plurality of blades 21 rotates about its respective pitchaxis P according to an airfoil schedule defined by the slot 130 to whichit is coupled by linkage arm 122. By defining different airfoil pitchschedules, the rotation of the blades 21 may be controlled independentlyof each other. Therefore, for example, if alternating blades 21 arerotated according to different airfoil pitch scheduling, conflictthrough flat pitch may be avoided. In addition, the pitch schedule maybe adjusted to improve performance of the blade 21. In certainembodiments, improved performance of the blade 21 via different airfoilpitch scheduling may reduce undesired acoustics, or mitigate theproduction of undesired acoustics, from the blades 21 during rotation atone or more operational modes of the engine 10, or during one or moreoperational modes of an aircraft to which the engine 10 is attached(e.g., takeoff, climb, cruise, approach, etc.).

The airfoil pitch schedules may depend, for example, on whether theaircraft is in a normal flight phase, a flat pitch transition phase, ora reverse thrust configuration. For example, the variable pitch rotorassembly 20 may be configured for normal flight phase when the blades 21have a pitch of greater than 8°. In addition, when the blades 21 arewithin 8° of flat pitch (i.e., between −8° and 8°), the variable pitchrotor assembly 20 may be operating in a flat pitch transition phase. Theblades 21 may be in a reverse thrust phase when angled at −8° or less.One skilled in the art will appreciate that these ranges are used onlyfor the purpose of explanation, and that phases and airfoil schedulesmay be defined in a variety of other ways to improve performance of thevariable pitch rotor assembly 20 and engine 10.

In an example embodiment, the plurality of blades 21 rotate according todifferent pitch schedules in order to avoid conflict as the blades 21rotate through flat pitch. More specifically, as shown in FIG. 2, afirst set of blades 134 may rotate according to a first airfoil pitchschedule, and an alternating, second set of blades 136 may rotateaccording to a second airfoil pitch schedule. The first and secondairfoil pitch schedule may be the same for a first phase of rotation,which may correspond to normal flight operation, but the pitch schedulesmay deviate from one another as the blades 21 enter flat pitch. Forexample, as soon as the pitch of the plurality of blades 21 reach within8° of flat pitch, the rotational speed of the first set of blades 134may increase while the rotational speed of the second set of blades 136may decrease. In this manner, the first set of blades 134 may passthrough flat pitch sequentially ahead of the second set of blades 136,thus avoiding contact through flat pitch. After all blades 21 havepassed through flat pitch and begin to generate reverse thrust, thefirst and second airfoil pitch schedules may once again synchronize witheach other so that all blades 21 rotate in unison. Alternatively,however, the airfoil schedules may remain offset in order to ensurereverse thrust is achieved without choking the air going to the core 16of the engine 10, or to achieve other performance improvements.

One skilled in the art will appreciate that the airfoil pitch schedulesdiscussed above are only exemplary, and that any other airfoil pitchschedule or schedules may be used as needed for performance. Forexample, more than two airfoil pitch schedules may be used. Indeed,every blade 21 could rotate according to its own pitch schedule. Allsuch variations are contemplated as within the scope of the presentdisclosure.

Now referring to FIGS. 6-7, a schematic representation of thedisplacement of the sliding member 132 is shown. This representationdepicts two adjacent blades 21 rotating according to airfoil schedulesdefined by scheduling slots 130 in scheduling ring 120. In theillustrated embodiment, each blade 21 is centered about respective pitchaxis P, where it is rotatably coupled to disk 42. Each linkage arm 122is schematically represented by dotted line 122 and rotates a fixedradial distance about its respective pitch axis P. Sliding member 132 isrotatably connected to linkage arm 122 and is slidably coupled toscheduling slot 130.

As shown in the figures, as scheduling ring 120 rotates relative to disk42, the scheduling slots 130 are generally translated in the directionindicated by arrow 140. For each angular position of the scheduling ring120, the angular position of each blade 21 may be varied according tothe shape of its respective scheduling slot 130, such as a firstscheduling slot 142 and a second scheduling slot 143. In variousembodiments, the first scheduling slot 142 defines a different contourfrom the second scheduling slot 143, such that each scheduling slot 142,144 rotates the blade 21 to a different position, or at a different rateof change, relative to one another. For example, referring specificallyto FIG. 6, some scheduling slots 130 may be entirely linear in thevertical direction (e.g., the first scheduling slot 142 defining alinear scheduling slot). By contrast, some scheduling slots 130 may benon-linear (e.g., the second scheduling slot 143 defining a non-linearscheduling slot), for example, by having one or more linear portions 146and one or more non-linear portions 147. In other example embodiments,the scheduling slots 130 may be bent, curved, serpentine, or any othersuitable shape.

Notably, when the scheduling ring 120 is rotated at a constant velocity,a linkage arm 122 connected to the entirely linear scheduling slot 142will have a constant rotational speed about pitch axis P. By contrast,the rotational speed of a linkage arm 122 connected to a non-linear slotwill vary according to the shape of its respective scheduling slot 130.In this manner, by alternately shaping each scheduling slot 130,alternating blades 21 may rotate into flat pitch at different times,such that blade 21 contact will not occur through flat pitch. Inaddition, adjacent scheduling slots 130 may have a similar profilethroughout the blade 21 angle range, such that the blades 21 rotate inunison throughout their range with the exception of the point where theyenter flat pitch.

One skilled in the art will appreciate that the above-describedmechanism for actuating the rotation of the rotor blades is only oneexemplary mechanism for achieving asynchronous rotor blade pitching.Other mechanisms will be evident to a skilled artisan based on thepresent disclosure. Any such variations or modifications arecontemplated as within the scope of the present disclosure.

The above-described embodiments facilitate thrust vector adjustment,including thrust reverse, thrust magnitude change and/or thrustdirection change along the longitudinal direction, for a variable pitchrotor assembly 20 with the blade 21 solidity greater than one without aneed for a heavy thrust reverse mechanism. Particularly, embodiments ofthe pitch change mechanism shown and described herein allows for atleast two-phase asynchronous blade 21 pitching, such that each blade 21rotates on a different schedule through flat pitch and/or reverseallowing the blades 21 to pass each other without contact. For example,the pitch change mechanism can rotate six out of twelve blades 21 on adifferent schedule through reverse, thus allowing reverse thrust to beachieved without contact between the blades 21 as they pass through flatpitch. All blades 21 may rotate on the same schedule throughout theentire flight envelope with the exception of the reverse condition.Benefits of asynchronous blade 21 pitching include improvements inengine efficiency and specific fuel consumption. Installation is alsosimplified as compared to prior designs, fan operability is improved,and stall margin is increased. Other advantages will be apparent tothose of skill in the art.

Referring back to FIG. 1, and further in conjunction with FIGS. 8-14, incertain embodiments, the vane assembly 30 includes a plurality of vaneairfoils 31 arranged in a spaced apart manner. Referring briefly to FIG.8, an exemplary airfoil 31 is provided graphically depicting how variousparameters such as camber and stagger angle are defined with respect tothe airfoil, such as the blade 21 (FIG. 1) or the vane 31 (FIG. 1). Anairfoil meanline is described as a line that bisects the airfoilthickness (or is equidistant from the suction surface and pressuresurface) at all locations. The meanline intersects the airfoil at aleading edge (LE) and a trailing edge (TE). The camber of an airfoil isdefined as the angle change between the tangent to the airfoil meanlineat the leading edge and the tangent to the angle meanline at thetrailing edge. The stagger angle is defined as the angle the chord linemakes with the centerline axis (e.g., reference line 44). Reference line44 is parallel to axis 11, and reference line 55 is orthogonal toreference line 44.

Referring generally to FIG. 1 and FIGS. 15-16, a vane characteristicsactuation assembly 148 in accordance with an exemplary embodiment of thepresent disclosure is depicted. In certain embodiments, the engine 10includes a pitch actuation assembly 48 at the rotor assembly 20 (e.g.,such as depicted and described in regard to FIGS. 2-7) and a vanecharacteristics change assembly 148 at the vane assembly 30 (e.g., suchas depicted and described in regard to FIG. 1 and FIGS. 15-16) todesirably control thrust output, thrust vector, rotor speed, acousticnoise, or generally allow for constant or substantially constant speedor operation of the core engine 40 while desirably adjusting magnitudeand/or direction of thrust output.

As mentioned above, one or more of the plurality of vanes 31 isrotatable about a vane pitch axis (e.g., vane pitch axis 90 in FIG. 1,FIGS. 15-16). The vane characteristics actuation assembly 148 mayprovide to one or more of the vanes 31 collective, independent, organged (i.e., a first set of vanes differently and/or independentlyoperable from a second set of vanes, such as depicted and describedherein) adjustment of the orientation or airfoil characteristics of thevane 31 about the vane pitch axis of each respective vane 31. Suchindependent or collective adjustment of pitch angle of the vane 31 aboutthe vane pitch axis may be utilized according to one or more methodsfurther described herein, such as one or more methods for attenuatingundesired acoustic noise, for producing a desired thrust vector, and/orfor producing a desired thrust load.

FIGS. 9-13 each include illustrations of radial sections of the engine10 taken through stages of axial flow airfoils and nearby aircraftsurfaces, and are typically referred to as “roll-out-views”, such as aprojection of blades about circumference onto a plane. These views aregenerated by sectioning airfoil stages and aircraft surfaces at a fixedradial dimension measured radially from longitudinal axis 11 andreference dimension R in FIG. 1. When blades 21 and vanes 31 ofrespective rotor assembly 20 and vane assembly 30 are sectioned atreference dimension R, corresponding blade 21 and vanes 31 aregenerated. Then the blades 21 and vanes 31 are unrolled or ‘rolled-out’to view the sections in two-dimensional space while maintaining thecircumferential and axial relationships between the airfoil stages andany nearby aircraft surfaces. Reference dimension E for the axialspacing between blades 21 and vanes 31. This allows the rotor assembly20 and the vane assembly 30 in FIGS. 9-13 to be described in twodimensions. An axial dimension, parallel to the longitudinal axis 11 andgenerally aligned with the direction Z of the moving working fluid shownin FIG. 1, and a ‘rolled-out’ or flattened circumferential dimension X,orthogonal to the axial dimension.

FIG. 9 illustrates a cross-sectional “roll-out view” of rotor assembly20 which as depicted includes twelve blades 21. Each blade 21 isindividually labeled with lower case letters o through z, with the blade21 labeled o repeating at the end of the sequence to highlight theactual circumferential nature of rotor assembly 20. Each blade 21 has ablade leading edge 244. A line positioned in the circumferentialdirection X through each blade leading edge 244 defines a rotor plane24. Each blade 21 is spaced apart from one another and is locatedaxially at the rotor plane 24.

Similar to the rotor assembly 20, the vane assembly 30 depicted in FIG.9 has ten vanes 31, individually labeled a through j, each with a vaneleading edge 333. A line positioned in the circumferential directionthrough each vane leading edge 333 defines a stator plane 34. In FIG. 9,each vane 31 in the vane assembly 30 is identical in size, shape, andconfiguration, and is evenly spaced circumferentially from each other(i.e., along reference dimension P) and evenly spaced axially from therotor plane 24 (i.e., in regard to reference dimension E). A nominal,evenly distributed circumferential spacing Q, between vanes 31 can bedefined by the following equation using the radial height of thereference dimension R, and the number of vanes 31, N, in vane assembly30:

Q=R*2*π/N

The engine 10 may include a controller configured to adjust the positionof one or more vanes 31, the blade pitch angle of the plurality ofblades 21 at the rotor assembly 20, and/or the rotor plane 24 of therotor assembly 20 relative to the plurality of blades 21 of the rotorassembly 20. In certain embodiments, the pitch angle at pitch axis(e.g., vane pitch axis 90 in FIG. 1), the longitudinal or axial spacingof a respective vane leading edge 333 to the rotor plane 24, and/or thecircumferential spacing of two or more vanes 31 along referencedimension Q is adjusted to improve the acoustic signature of the engine10 relative to various operational conditions of the engine 10 and/orthe aircraft (e.g., angle of attack). Exemplary embodiments ofadjustments or positioning of the vane assembly 30 relative to the rotorassembly 20 are further provided in regard to FIGS. 10-12. In each ofthese figures, the rotor assembly 20 and vane assembly 30 are locatedaxially forward of a wing of an aircraft. Additionally, an exemplaryembodiment of an aircraft surface 1160 is represented as two wingsections 1161, 1162. Note that two wing sections are present in each“roll-out view,” because the radial section that generates theseinstalled views cuts through the wing of an aircraft in twocircumferential locations. For the non-uniform vanes 31 in all of theFigures which follow, this dashed and solid line depiction method isused to refer to exemplary embodiments of nominal and non-nominal vanes31 respectively.

To minimize the acoustic signature, it is desirable to have theaerodynamic loading of the vane leading edges 333 to all be similar andbe generally not highly loaded. To maximize the efficiency and minimizethe acoustic signature of the rotor assembly 20, a desired goal would beto minimize the variation in static pressure circumferentially along therotor assembly 20. To maximize the performance of the vane assembly 30,another goal would be to have neither the aerodynamic loadings of thevane leading edges 333 nor the vane suction 35 and pressure surface 36diffusion rates lead to separation of the flow.

To maximize the performance of the aircraft surface, depicted in theseexemplary embodiments as a wing sections 1161 and 1162, one goal may beto keep the wing loading distribution as similar to the loadingdistribution the wing was designed for in isolation from the engine 10,thus maintaining its desired design characteristics. The goal ofmaintaining the aircraft surface 1160 performance as designed for inisolation from the engine 10 applies for aircraft surfaces that may benon-wing, including, for example, fuselages, pylons, and the like.Furthermore, to maximize the performance of the overall aircraft andengine 10, one of the goals would be to leave the lowest levels ofresultant swirl in the downstream wake. As described herein, thenon-uniform characteristics of the vanes 31 is adjusted based on one ormore of these desired goals during operation of the engine 10 andaircraft.

This optimal performance can be accomplished in part by developingnon-uniform vane exit flow angles, shown in FIG. 10 as angles Y and Z,to minimize interaction penalties of the engine installation and toreduce the acoustic signature. The first exemplary embodiment of this isshown in FIG. 10, where each pair of vanes 31 in the vane assembly 30are evenly spaced circumferentially from one another and evenly spacedaxially from the rotor plane 24. However, the nominal (without pitchchange) stagger angle and camber of the vanes 31 in FIG. 10 vary toprovide optimal exit flow angles into the aircraft surface downstream ofthe vane assembly 30, such as depicted in regard to reference vanes 31labeled b through e, and g through i.

FIG. 11 shows another exemplary embodiment of vane assembly 30 providingflow complimentary to aircraft surface 1160. In FIG. 11, vanes 31 andrelated vanes 31 in vane assembly 30 are not evenly spacedcircumferentially from each other, nor are they evenly spaced axiallyfrom the rotor plane 24. The degree of non-uniformity may vary along thespan of a vane. Two vanes 31 are spaced axially forward of the statorplane 34, reference dimensions F and G, allowing the vane assembly 30 tomerge axially with the aircraft surface 1160. For instance, the aircraftsurface 1160 may at least partially include or define at least one ofthe vanes 31 of the vane assembly 30. The nominal (without pitch change)stagger angle and camber angle of the vanes 31 vary to provide optimalexit flow angles into the wing sections 1161 and 1162, as shown in vanes31 labeled d through i.

FIG. 12 is similar to FIG. 11, but depicts the removal of two vanes 31adjacent to wing section 1161. This exemplary embodiment allows thevanes 31 to be evenly spaced axially from the rotor plane 24 and allowsthe wing section to merge axially with the vane assembly 30.

Although the location of the rotor assembly 20 and vane assembly 30 ineach of the foregoing exemplary embodiments was axially forward of theaircraft surface 1160, it is foreseen that the propulsion system 70could be located aft of the aircraft surface 1160. In these instances,the prior enumerated goals for optimal installed performance areunchanged. It is desirable that the propulsion system has suitable rotorassembly 20 circumferential pressure variations, vane leading edge 333aerodynamic loadings, and vane pressure surface 35 and suction surface36 diffusion rates. This is accomplished in part by varying the size,shape, and configuration of each vane 31 and related vane 31 in the vaneassembly 30 alone or in combination with changing the vane 31 pitchangles. For these embodiments, additional emphasis may be placed onassuring the combined propulsion system 70 and aircraft leave the lowestlevels of resultant swirl in the downstream wake.

The exemplary embodiment of the rotor assembly 20 and vane assembly 30in FIG. 9 is designed for a receiving a constant swirl angle, referenceangle A, into vanes 31 along the stator plane 34. However, as theaircraft angle of attack is varied the vanes move to off designconditions and the swirl angle into the vane assembly 30 will varyaround the stator plane 34. Therefore, to keep the aerodynamic loadingon the vane leading edges 33 roughly consistent along the stator plane34, a variable pitch system that would rotate either each vane 31 orgroup of vanes 31 a different amount is desirable. Such a pitch changecan be accomplished by rotating a vane 31 in a solid body rotation alongany axis, including, for example, the axis along the centroid of vane 31or an axis along the vane leading edge 333. The desire for similaraerodynamic loading on the vane leading edges 33 is in part driven bythe desire to keep the acoustic signature of the engine 10 low. Vanes 31with high leading edge loadings tend to be more effective acousticradiators of the noise created from the gust of the upstream rotorassembly 20. The exemplary embodiment of the rotor assembly 20 and vaneassembly 30 in FIG. 13 describes this desired variation in vane 31 viachanges in pitch angles of one or more vanes 31, such as via the vanepitch change mechanism 148 further described herein. For ease ofexplanation, we define the chord line angle of vanes at the design pointas stagger and hence variations between vanes at the design point asstagger variations. As the engine 10 moves to different operatingconditions, or as the aircraft to which the engine is attached moves todifferent operating conditions (e.g., takeoff, climb, cruise, approach,landing, etc.), vanes 31 may rotate around the pitch axis 90 referred toas pitch change (or changes in pitch angle) of the vanes 31. Variationsin vane chord angles that result from these solid body rotations arereferred to as pitch angle variations.

In FIG. 13, each vane 31 in the vane assembly 30 is identical in size,shape, and configuration, and are evenly spaced circumferentially fromeach other and evenly spaced axially from the rotor plane 24. However,the pitch angles of the vanes 31 in FIG. 13 vary as they represent achange in the vane 31 pitch actuation to accommodate varying inputswirl, reference different input swirl angles A and B, into stator plane34 caused in part by changes in aircraft angle of attack. As desired,this provides similar aerodynamic loading on the vane leading edges 33to keep the acoustic signature of the engine 10 low, such as within oneor more ranges further described herein. This similar loading can beaccomplished by independently changing pitch angle for individual vanes31 via the vane characteristics change mechanism 148, or by changingpitch angles similarly for groups of vanes 31 suitable for ganging. Thevanes 31 could rotate in pitch about any point in space, but it may bedesirable to maintain the original leading edge 333 circumferentialspacing and rotate the vanes 31 around a point at or near their leadingedge 333. This is shown in FIG. 13 using vanes 31 labeled c, d, f, andg, where the nominal staggered vanes 31 are depicted in dashed lines andthe rotated (or pitched) vanes 31 are depicted as solid lines.

As shown by way of example in FIG. 14, it may be desirable that eitheror both of the sets of blades 21 and vanes 31 incorporate a pitch orairfoil characteristics change mechanism (e.g., blade pitch actuationassembly 48 in FIGS. 2-7, vane characteristics change mechanism 148 inFIGS. 15-16) such that the blades and vanes can be rotated with respectto an axis of pitch rotation either independently or in conjunction withone another. Such pitch change can be utilized to vary thrust and/orswirl effects under various operating conditions, including providingthrust reversing, acoustic noise attenuation, or desired thrust vector,which may be useful in certain operating conditions of the engine 10and/or aircraft.

The vane system 30, as suitable for a given variation of input swirl andaircraft surface 1160 installation, has non-uniform characteristics orparameters of vanes with respect to one another selected either singlyor in combination from those which follow. A delta in stagger anglebetween neighboring vanes 31 according to one embodiment of greater thanor equal to about 2 degrees can be employed, and according to anotherembodiment between about 3 degrees and about 20 degrees. A delta incamber angle between neighboring vanes 31 and related vanes 31 accordingto one embodiment of greater than or equal to about 2 degrees can beemployed, and according to another embodiment between about 3 degreesand about 15 degrees. A circumferential spacing Q at a given referencedimension R, between neighboring vanes 31 and related vanes 31, for vane31 counts N from about 5 to about 20, from about 10% to about 400% ofthe nominal, even circumferential spacing can be employed. An axialspacing from the rotor plane 24 to vanes 31 and related vanes 31 up toabout 400% of the radial height H, of the vane 31 can also be employed.

The non-uniform characteristic may be attributed to a portion of thespan of the vanes, or to substantially all of the span of the vanes. Incertain embodiments, at least a portion, or all, of the plurality ofvanes 31 of the vane assembly 30 may include the vane characteristicsactuation mechanism 148, in which the vane characteristics actuationmechanism is configured to adjust at least a pitch axis and/or axialspacing such as described herein.

Still various embodiments of the vane assembly 30 provided herein mayinclude at least one vane defining a pylon or aircraft surface (e.g.,aircraft surface 1160). It should be appreciated that vane pitch anglechanges may desirably alter thrust direction to or away from the pylonsurface, such as described herein, to attenuate generation of undesirednoise. In certain embodiments, one or more aircraft surfaces, such asthe pylon, may include pitch change mechanisms, flaps, or actuatorsconfigured to perform substantially similarly as one or more vanesdepicted and described herein.

Referring now to FIGS. 17-21, exemplary depictions of adjustments,actuations, or other changes in pitch at the blades 21 and/or vanes 31are provided. FIGS. 17-21 provide a radial view of an airfoil profilesuch as described in regard to FIGS. 1-16, corresponding to a radiallocation for which a contribution to reverse thrust is desired, such asthe outer span of the blade 21. In regard to FIGS. 17-21, closing theblade 21, such as changing the pitch of the blade 21 toward a closedposition, is represented by a clockwise rotation (e.g., depicted viaarrows CW) of the airfoil about its respective pitch axis 91. Closingthe vane 31, such as changing the pitch of the vane 31 toward a closedposition, is represented by a counter-clockwise rotation (e.g., oppositeof arrows CW) of the airfoil about its respective pitch axis 90. In thevector diagrams depicted in FIGS. 17-21, subscript 1 (e.g., V₁, W₁)refers to an airflow condition at a first station forward of the rotorassembly 20 (e.g., proximate to forward end 98). Subscript 2 (e.g., V₂,W₂) refers to an airflow condition at a second station between the rotorassembly 20 and the vane assembly 30. Subscript 3 (e.g., V₃, W₃) refersto an airflow condition at a third station aft of the vane assembly 30(e.g., proximate to aft end 99). V₁, V₂, V₃ each refer to absolutevelocity at their respective airflow stations. W₁, W₂, W₃ each refer tovelocity relative to a rotating frame of reference of the rotor assembly20 at their respective airflow stations. U indicates magnitude anddirection of the speed of the rotor blade 21 corresponding to therotational speed and radial location. Axial and tangential velocitycomponents are indicated by the vertical and horizontal components,respectively, of the vectors. Positive tangential velocity componentsare in the direction of blade speed vector U. The change in tangentialvelocity that occurs as the flow travels through rotor assembly 20,indicated by ΔV_(t), indicates the level or magnitude of loading on therotor.

It should be appreciated that as the rotor assembly 20 rotates about thelongitudinal axis 11 of engine 10, it imparts tangential momentum, orswirl, to the flow, such that the flow exiting the rotor assembly hasgreater tangential velocity than the flow entering it. If flow entersthe rotor assembly 20 with zero or substantially zero tangentialvelocity then the flow exits the rotor with a positive tangentialvelocity. The flow exiting the rotor assembly 20 may have a positivetangential velocity component, a zero or substantially zero tangentialvelocity component, or a negative tangential velocity component based onthe flow entering the rotor assembly 20, the blade pitch angle aboutpitch axis 91, or speed U of the rotor blade 21, or combinationsthereof. Residual swirl from forward thrust and reverse thrust operation(i.e., non-zero tangential velocity component in the flow exiting thepropulsion system) does not contribute substantially to the thrustcapability of the system.

FIG. 17 illustrates the engine 10 during forward thrust operation. Inthe embodiment depicted in FIG. 17, the rotor assembly 20 and the vaneassembly 30 are each at or near their exemplary design point or forwardthrust position (e.g., nominal position). As the air flows from thefirst station to the second station, the rotor assembly 20 impartstangential momentum to the air such as to increase the tangentialcomponent of the absolute velocity (V) in the direction of rotation.From the second station to the third station, the vane assembly 30removes a substantial portion, or all, of tangential momentum from theflow from the rotor assembly 20, resulting in an exit velocity from thevane 31 at the third station that has less tangential velocitycomponent. As such, the lower exit tangential velocity component maydefine, at least in part, a desirably efficient generation of thrust viareduced waste in kinetic energy. Stated differently, the desirablyefficient generation of thrust directs kinetic energy generally alongthe axial or longitudinal direction rather than along a non-axialdirection.

FIG. 18 illustrates the engine 10 during an exemplary thrust reverseoperation mode. The vane 31 is positioned substantially at or near itsdesign point pitch angle, such as depicted in regard to FIG. 17. Forreverse thrust operation, the blade 21 is closed, such as via rotationalong blade pitch axis 91 in the CW direction, so that the leading edge244 (i.e., the thick end relative to a thinner trailing edge 245) is aftof the trailing edge 245. As the rotor assembly 20 rotates about thelongitudinal axis 11 of engine 10 with the blade 21 in closed position,the rotor assembly 20 induces flow from the second station to the firststation. As a result, the vane 31 imparts a tangential velocitycomponent opposite to the direction of rotation of the rotor assembly 20(i.e., counter-swirl) as the flow progresses from the third station tothe second station. Absolute velocity V₂ has a negative tangentialvelocity component, and the effect of the rotor assembly 20 is to resultin absolute velocity V₁ with reduced or eliminated negative tangentialvelocity component. This results in increased rotor loading andincreased reverse thrust output relative to a system without astationary vane row. With increased rotor loading, the change in angleof the relative velocity from W₂ to W₁ increases due to a highernegative tangential component of W₂. With the increase in loading on therotor, the departing reverse thrust relative velocity W₁ may have ahigher negative tangential component and less swirl component in thedeparting absolute velocity W₁ than a system without a stationary vanerow.

Referring now to FIG. 20, another thrust reverse operation mode isdepicted. In FIG. 20, the vane 31 is closed (i.e., rotated CCW along thevane pitch axis 90) relative to its design point pitch angle (e.g., suchas the vane depicted in FIG. 17). Relative to the example depicted inFIG. 18, closing the vane 31 increases the counter-swirl entering therotor blade 21, such as to increase reverse thrust output via increasedrotor loading at the blade 21. The increased rotor loading may increasethe negative tangential component of the relative velocity W₁ departingthe blade 21 and provide a reduced exit swirl. If the increased loadingon the rotor blade 21 produces excess reverse thrust or exit swirl, therotor blade 21 may be opened (i.e., rotated CCW along the blade pitchaxis 91) to enhance the desired effect.

FIG. 21 depicts the vane 31 opened (i.e., rotated CW along the vanepitch axis 90) relative to its design point pitch angle. In certaininstances, methods or operations may desirably reduce or spoil thrustreverse such as depicted in FIG. 21, such as via opening the vane pitchangle to generate a resultant tangential velocity component at the firststation. Desirably reducing or spoiling thrust reverse via altering thevane pitch angle allows for reducing thrust without changing rotor bladepitch 91 angle or rotational speed at the rotor assembly 20. As such,reverse thrust output may be desirably altered without altering coreengine speed or output, or while maintaining substantially constant coreengine speed, such as described herein. It should be appreciated thatspoiling reverse thrust such as described herein may mitigate risks ordamage related to core engine or rotor assembly overspeed, changes intorque output at the core engine or rotor assembly, or other transientoperations at the core engine or rotor assembly. However, it shouldfurther be appreciated that alteration or adjustment of vane angle for adesired thrust output may be performed with changes at the core engineand/or rotor assembly.

Referring now to FIG. 19, another thrust reverse operation mode isdepicted. The vane 31 is opened (i.e., the vane 31 is rotated CW alongthe vane pitch axis 90) relative to its design point, or nominal, pitchangle such as to reduce the negative tangential velocity component atthe second station relative to the thrust reverse mode depicted inregard to FIG. 18. The reduced tangential velocity opposite of therotation of the blades 21 (i.e., counter-swirl) tends to unload therotor assembly 20. However, the pitch angle of the blade 21 is closed(i.e., rotated CW along the blade pitch axis 91) an additional amount toat least partially recover the rotor loading and impart a change inabsolute tangential velocity from the second station to the firststation, such as described in regard to FIG. 18. Reverse thrust outputmay be reduced or spoiled by way of opening the vanes 31 whilemaintaining a substantially constant operation of the core engine 40and/or without changing blade pitch angle or rotational speed.

In various embodiments, the blade 21 is formed such that the staggerangle varies from hub to tip to accommodate the desired flow vectors andloading distribution along the blade span. The varied stagger angle mayallow for forward thrust output from a lower span of the blade andreverse thrust output from an upper span of the blade. In oneembodiment, the blade is configured to generate forward flow below 50%span. In still another embodiment, the blade is configured to generatereverse flow at or above 50% span. In still various embodiments, thevane is configured to change angle up to 15 degrees open and up to 15degrees closed from the design point to adjust reverse thrust such asdescribed above. In another embodiment, the vane is configured to changeup to 10 degrees open and up to 10 degrees closed from the design pointto adjust reverse thrust such as described above. In yet anotherembodiment, the vane is configured to change angle up to 5 degrees openand up to 5 degrees closed to adjust reverse thrust such as describedabove.

Referring now to FIG. 22, a method for adjusting thrust vector for anunducted rotor engine is provided, (hereinafter, “method 1000”).Embodiments of the method 1000 provided herein may particularly providefor altering or adjusting thrust direction and magnitude for an unductedrotor engine. Certain embodiments provide for control and generation ofreverse thrust for an unducted rotor engine. Still certain embodimentsprovide for control and generation of thrust and desirably altering atangential velocity component of the flow exiting the rotor assembly.Embodiments of the method 1000 may be applied with gas turbine engineswith articulatable fan or propeller rotor pitch axis and vane assemblypitch axis, such as depicted in all or part of the embodiments of theengine 10 provided in regard to FIGS. 1-21. Steps of the method 1000 maybe store and executed via a controller, such as controller 210 depictedin FIG. 1, or one or more controllers depicted and described in FIGS.23-29. However, it should be appreciated that the method 1000 may beexecuted with other configurations of unducted rotor engine.

The method 1000 includes at 1010 generating, via rotation of a rotorassembly, a flow of air from a first station forward of the rotorassembly to a second station aft of the rotor assembly. In a particularembodiment, the rotor assembly generates the flow of air from forward ofthe rotor assembly to aft of the vane assembly. In certain embodiments,the method 1000 further includes generating an aft axial velocitycomponent of the flow of air from aft of the vane assembly to forward ofthe rotor assembly at least by closing a blade pitch angle at one ormore blades at the rotor assembly. At 1020, the method 1000 includesgenerating a positive tangential velocity component of the flow of airvia the rotor assembly. It should be appreciated that the positivetangential velocity component is along a first direction correspondingto the direction of rotation of the rotor assembly. A negativetangential velocity component is along a second direction opposite ofthe first direction.

In certain embodiments, the method 1000 includes at 1030 increasingloading at the blades of the rotor assembly. In one embodiment,increasing loading at the blades of the rotor assembly includes closingthe blade pitch angle when generating reverse thrust via a flow of airfrom aft of the rotor assembly to forward of the rotor assembly, such asdescribed in regard to FIG. 18. In another embodiment, increasingloading at the blades of the rotor assembly includes closing a vanepitch angle at the vane assembly. In one embodiment, increasing loadingat the rotor assembly includes at 1032 closing the pitch angle of thevane at the vane assembly and the blade at the rotor assembly. Incertain embodiments, the method 1000 includes at 1034 reducing thenegative tangential velocity component of the flow of air from the vaneassembly at least by opening the vane pitch angle.

In still certain embodiments, the method 1000 includes at 1040 adjustingthe blade pitch angle at one or more blades of the rotor assembly toposition a blade leading edge aft of a blade trailing edge, such asdepicted and described in regard to FIGS. 17-19. In particularembodiments, rotating the rotor assembly after adjusting the blade pitchangle generates the positive tangential velocity component of fluid fromthe rotor assembly.

In still various embodiments, the method 1000 includes at 1050 adjustingthe vane pitch angle at one or more vanes of the vane assembly. In oneembodiment, adjusting the vane pitch angle includes reducing loading atthe rotor assembly. In certain embodiments, the method 1000 includes at1052 changing absolute tangential velocity relative to aft and forwardof the rotor assembly. Changing absolute tangential velocity is based atleast on adjusting blade pitch angle relative to vane pitch angle.

In certain embodiments, the method 1000 includes at 1070 rotating thevane at the vane assembly co-directional to a direction of rotation ofthe blade pitch angle at one or more blades of the rotor assembly. Inone embodiment, the method 1000 includes at 1072 closing the vane pitchangle at one or more vanes of the vane assembly. In a particularembodiment, closing the vane pitch angle at one or more vanes of thevane assembly includes reducing a negative tangential velocity componentat the vane assembly. In still another embodiment, closing the vanepitch angle at one or more vanes of the vane assembly includesincreasing counter-swirl of the flow of air from the vane assembly, suchas described and depicted in regard to FIGS. 20-21.

Referring now to FIGS. 23-29, diagrams outlining steps for operation ofan unducted rotor engine are provided. The methods and diagrams providedherein may be utilized with various embodiments of a single unductedrotor engine, such as engine 10 depicted and described herein. However,it should be appreciated that the methods provided herein may beutilized to control engines generally including one or more of a rotorblade pitch change mechanism, a vane pitch change mechanism, a rotorangle of attack change mechanism, or combinations thereof.

Conventional turbofan engines generally control engine thrust bymeasuring corrected fan speed or overall engine pressure ratio andcorrelating one or both measurements to desired engine thrust based onan aircraft flight condition. However, methods for controlling andoperating an unducted single rotor engine, such as depicted anddescribed herein, include adjusting, to generate and adjust thrustoutput, a rotational speed (e.g., mechanical speed, corrected speed,etc.) of the rotor assembly (e.g., rotor assembly 20), a rotor bladepitch (e.g., at axis 91) at the rotor assembly, a torque (e.g., torqueon a fan or propeller shaft), engine pressure ratio (e.g., P56/P2), orcore engine pressure ratio (e.g., P56/P25). The method includesgenerating or adjusting thrust output based at least on a performancemap, curve, table, or other reference position or function of the rotorassembly as a function of rotor blade pitch (e.g., at axis 91). Invarious embodiments, the method further includes generating or adjustingthrust output based on rotational speed of the rotor assembly or aflight condition (e.g., takeoff, climb, cruise, approach, landing, etc.,or one or more air conditions related thereto, including air speed,pressure, temperature, density, humidity, or other environmentalcondition).

In certain embodiments, the method includes determining or adjustingengine thrust output based at least on an engine cycle model, such asthe power management and engine cycle model block (power managementblock) depicted in FIGS. 23-29. In one embodiment, a controller 1600includes one or more, or multiple, single-input, single-output (SISO)loops or a combination of SISO and multi-input, multi-output (MIMO)loops to provide dynamic coordination or adjustment between blade pitchchanges (e.g., blade pitch 91 at one or more blades 21), vane pitchchanges (e.g., vane pitch 90 at one or more vanes 31), rotor planechanges (e.g., rotor plane 34), core engine speed changes (e.g., coreengine 40), electric machine load changes, or combinations thereof.

Referring to FIG. 23, operations or method steps executed at thecontroller 1600 may include receiving or obtaining, at the powermanagement block in the control logic, a throttle input. In variousembodiments, the schematic controller 1600 depicted in regard to FIG. 21is a sensor-based controller that infers a desired thrust output basedon a lookup table, chart, schedule, or other reference. The throttleinput is mapped to one or more of a desired rotor speed, a desiredtorque, a desired thrust output, a desired pressure ratio across theengine and/or rotor assembly, while adjusting to or otherwise accountingfor environmental conditions related to an aircraft state at thecorresponding moment or period of time (e.g., air speed, pressure,temperature, density, humidity, altitude, etc.). In a first controlloop, the engine cycle model or power management block outputs a signalindicative of a desired thrust output. In certain embodiments, thesignal is a commanded rotor blade pitch angle and/or commanded vanepitch angle, such as described in regard to rotor assembly 20 and/orvane assembly 30 herein. A difference between a commanded pitch angle,measured pitch angle (i.e., pitch error), is received at a controller(e.g., Control) configured to control, adjust, or otherwise articulatethe rotor blade pitch and/or vane pitch, such as via a respective rotorblade pitch change mechanism or vane pitch change mechanism. The engine10 adjusts the rotor blade pitch angle 91 based on the commandedadjustment. In various embodiments, a pitch sensor obtains, receives,measures, or otherwise acquires an actual rotor blade pitch angle at oneor more of the plurality of rotor blades (e.g., blade 21), at asynchronization ring (e.g., scheduling ring 120), or at one or moreslots (e.g., slot 130) at the rotor blade pitch change mechanism (e.g.,blade pitch change mechanism 48). The pitch sensor provides an outputsignal to a difference function, at which a difference or delta betweenthe commanded rotor blade pitch angle from the engine cycle model iscompared to the actual rotor blade pitch angle obtained from the pitchsensor.

In certain embodiments, the engine 10 adjusts the vane pitch angle 90based on the commanded adjustment. In various embodiments, a pitchsensor obtains, receives, measures, or otherwise acquires an actual vanepitch angle at one or more of the plurality of vanes (e.g., vane 31), ata synchronization ring, or at one or more slots at the vane pitch changemechanism (e.g., vane pitch change mechanism 148). The pitch sensorprovides an output signal to a difference function, at which adifference or delta between the commanded vane pitch angle from theengine cycle model is compared to the actual vane pitch angle obtainedfrom the pitch sensor.

Referring still to FIG. 23, the controller 1600 includes a secondcontrol loop, at which the engine cycle model outputs a commanded lowspool parameter (e.g., N₁-cmd, such as referring to a low speed spool5054 including the low speed compressor 5052 and the low speed turbine50 in FIG. 1). A difference (e.g., N₁ error), based at least on thecommanded low spool parameter and adjusted based at least on an actualN₁ parameter obtained or measured from an N₁ sensor (e.g., N₁ feedback),is provided to the controller (e.g., Control). The controller outputs anengine control signal, including, but not limited to, a commanded fuelflow to the engine (e.g., combustion section 4048 in FIG. 1). Thecommanded fuel flow to the engine, such as the combustion section,includes or one or more of a fuel flow rate, pressure, temperature, or avalve, orifice, manifold, area, or volume adjustment corresponding tothe commanded fuel flow, or other parameter that may affect an amount offuel provided to the combustion chamber for generating combustion gases.The actual N₁ parameter corresponds to the commanded N₁ parameter andmay include one or more of a low speed spool rotational speed (e.g.,mechanical speed, corrected speed, etc.) torque (e.g., propeller shafttorque), or pressure ratio (e.g., pressure ratio across one or morecompressors, or overall pressure ratio across the engine or coreengine).

Although FIG. 23 depicts a first control loop and a second control loopcontrolling rotor blade pitch 91 and rotational speed of the rotorassembly 20, it should be appreciated in other additional or alternativeembodiments, control for rotational speed of the rotor assembly 20 maybe replaced by, or augmented by, control of torque or engine pressureratio (e.g., pressure measured, by a sensor, downstream of the turbinesection, by a sensor, upstream of the compressor section over pressuremeasured, such as P56/P2) or core engine pressure ratio (e.g., pressuremeasured downstream of a turbine of the turbine section over pressuremeasured between a low pressure compressor and a high pressurecompressor, such as P56/P25), or core engine rotational speed.

Referring now to FIG. 24, in yet other embodiments, steps of the foroperating the engine may include providing a controller 1610 includingtwo or more control loops. Various embodiments of the controller 1610may be configured substantially similarly as depicted and described inregard to controller 1600 in regard to FIG. 23. In certain embodiments,the controller 1610 may include two or more control loops configured tocontrol an engine using at least two parameters indicative of thrust. Itshould be appreciated that in certain embodiments, the two or moreparameters indicative of thrust include a vane pitch angle 90 at a vaneassembly 30 positioned in aerodynamic relationship with the rotorassembly (e.g., the vane assembly 30 with adjustable vane pitch angle 90relative to the rotor assembly 20). In one embodiment, such as depictedin FIG. 24, the two or more closed control loops may include acombination of a fan or propeller speed and fan or propeller systemtorque, or a combination of fan or propeller speed and a pressure ratiosuch as described herein.

In still various embodiments, the controller 1610 may include three ormore control loops. For instance, a third loop may add another feedbackparameter based on sensed or calculated variables to manipulate ormodulate another effector such as other variable geometry (VG) (e.g.,one or more stator vanes, an inlet guide vane, bleed valves, etc.) orother mechanism for modulating power or airflow. Control of two or moreloops may be performed by a multi-input, multi-output (MIMO) controller,such as depicted in FIGS. 24-25. In still various embodiments, controlof two or more loops may be performed by several single-input,single-output (SISO) controllers, or combinations of MIMO and SISOcontrollers.

Referring now to FIG. 25, another embodiment of a controller 1700configured to execute steps of the method for operating a singleunducted rotor engine is provided. The methods and diagrams providedherein may be utilized with various embodiments of a single unductedrotor engine such as depicted and described herein. However, it shouldbe appreciated that the methods provided herein may be utilized tocontrol engines generally including one or more of a rotor blade pitchchange mechanism, a vane pitch change mechanism, a rotor angle of attackchange mechanism, or combinations thereof. The power management blockmay further output an aircraft power extraction signal indicative ofbleed air, electrical load, or other power extractions from one or moreengines for aircraft systems (e.g., thermal management, environmentalcontrol system, electrical or electronic systems, etc.).

Referring to FIGS. 26-29, the schematic controller 1800 depicted isconfigured as a model-based controller, such as depicted at FIG. 26,configured to calculate a desired output thrust based at least on anengine operating parameter including fuel flow, variable geometry (e.g.,vane pitch angle, blade pitch angle, compressor vane or bleedopening/closing, fuel flow controls, etc.), and flight condition(altitude, Mach number, ambient air temperature and/or pressure, one ormore aircraft loads, such as, but not limited to, aircraft electricalloads, thermal or environmental control system bleeds, or other aircraftbleeds and power extractions). Steps of the method for operation mayinclude receiving or obtaining at the controller 1800, or particularlyat the power management block of the control logic, a desired thrustoutput, such as from a throttle input. The power management blockdetermines or otherwise calculates commanded positions for variablegeometries (VGs) (e.g., variable vane angles, actuator positions, bleedvalve open/close positions, or other variable geometries) as well as thecommanded fan speed or corrected fan speed. The controller (e.g.,Control) receives the difference between the commanded thrust outputsignal and an actual or estimated thrust output signal from an enginemodel and tracking filter. The controller provides to the engine anoutput signal corresponding to one or more of a commanded fuel flow(e.g., flow rate, pressure, temperature, etc.), a rotor blade pitchangle, a vane pitch angle, a rotor plane angle at the rotor assembly, orone or more other variable geometries (e.g., Other VGs), such as, butnot limited to, a variable vane or bleed valve at a compressor section,a bypass valve or flow, a turbine nozzle area, a mid-fan inlet guidevane for a third-stream flowpath, booster or low pressure compressorvariable stator vanes, or third-stream variable nozzle, or one or moreactuator positions corresponding thereto. The engine receives the outputsignal from the controller and generates an actual thrust output basedon the output signal from the controller. One or more engine signalscorresponding to one or more engine sensors, such as a torque sensor,low spool speed measurement (e.g., N₁ speed), high spool speedmeasurement (e.g., N₂ speed), other spool speed measurements (e.g.,intermediate spools for 3-shaft engines, fan shaft speeds for gearedengine arrangements, etc.), a rotor blade pitch angle measurement, avane pitch angle measurement, a rotor plane position measurement, or oneor more actuator positions corresponding to the variable geometriesarticulated by the output signal, or an acoustic sensor (e.g.,microphone, vibration sensor, accelerometer, etc.) is provided from theengine and obtained by the engine model and tracking filter. A thrustfeedback signal is generated from the engine model and tracking filterbased on one or more engine signals corresponding to one or more enginesensors.

Referring to FIGS. 26-29, in certain embodiments, the engine, aircraft,system, or method may include a computing system 1800 including asensor-based controller, such as depicted in FIGS. 27-29, and amodel-based controller, such as depicted in FIG. 26. The computingsystem 1800 may be configured substantially similarly such as depictedand described in regard to a plurality of control devices such asdepicted and described in regard to FIGS. 23-25 (e.g., controller 1600,1610, 1700). In one embodiment, a model-based controller, such asdepicted in FIG. 26 or configured such as controller 1700 depicted inregard to FIG. 23, is utilized as a supervisory controller to provide atrim or adjustment to engine operation and performance, such asdescribed herein. In one instance, the model-based controller may alteror vary engine output thrust within a 7% margin (e.g., +/−3.5%) of a setcondition. In another instance, the model-based controller may alter orvary engine thrust output within a 5% margin (e.g., +/−2.5%) of a setcondition. In yet another instance, the model-based controller may alteror vary engine thrust output within a 2% margin (e.g., +/−1%) of a setcondition.

In various embodiments, such as an exemplary embodiment depicted in FIG.26, the computing system 1800 includes a model-based trim functionproviding one or more trims or adjustments to controller commands toimprove engine performance or operability based on a current operatingstate of the engine, aircraft, and environmental parameter. In variousembodiments, the model-based or state-aware scheme includes a parameterestimation algorithm, also known as a tracking filter, to update themodel to match actual engine characteristics or engine health, such asengine deterioration.

The sensor-based controller, such as depicted in FIGS. 27-29 orconfigured such as controller 1600, 1610 depicted and described inregard to FIGS. 23-24, may provide control and adjustment for changes inthrust output greater than the trim or adjustment levels of themodel-based controller. In various instances, the sensor-basedcontroller may provide for transient changes in engine operatingcondition, such as to and between two or more of ignition, idle,takeoff, climb, cruise, descent, approach, or thrust reverse. In otherinstances, the model-based controller may provide for adjustments duringsubstantially steady state engine operating condition.

Referring to FIGS. 23-29, embodiments of methods for operation of anunducted rotor engine may adjust, change, articulate, or actuate one ormore of a rotor blade pitch, a vane pitch, a rotor plane (e.g., via anAngle of Attack change mechanism), low spool speed, high spool speed, orcombinations thereof, such as via one or more controllers 1600, 1610,1700, 1800, or combinations thereof.

As such, a computing system for an unducted rotor engine is provided inwhich the computing system includes one or the other, or both, of asensor-based controller or a model-based controller such as describedherein. The sensor-based controller is configured to execute a first setof operations, such as those described herein, including obtaining afirst signal corresponding to a commanded low spool speed. obtaining asecond signal indicative of a pitch angle corresponding to thrust outputfrom the unducted rotor assembly and variable pitch vane assembly, andgenerating a pitch feedback signal corresponding to a commandedadjustment to the pitch angle based at least on one or both of avariable blade pitch angle or a variable vane pitch angle.

In various embodiments, the first set of operations executed by thesensor-based controller include obtaining a throttle input correspondingto one or more of a desired air speed of an aircraft, a desired thrustoutput, or a desired pressure ratio, generating the first signalcorresponding to the commanded low spool speed, and generating thesecond signal indicative of the pitch angle corresponding to thrustoutput at the rotor assembly. In a particular embodiment, the first setof operations further includes generating a low spool speed feedbacksignal corresponding to the commanded fuel flow.

In certain embodiments of the computing system including the model-basedcontroller and the sensor-based controller, the sensor-based controlleris configured to generate the pitch feedback signal during transientchanges in engine operating condition. Such as described herein,transient changes in engine operating condition may include conditionsto and between two or more of ignition, idle, takeoff, climb, cruise,descent, approach, or thrust reverse.

Additionally, or alternatively, the computing system includes themodel-based controlled configured to execute a second set of operationssuch as described herein, including obtaining a desired thrust outputvia a throttle input, determining, at least via a power managementblock, a commanded thrust output signal, receiving the commanded thrustoutput signal, and generating an output signal corresponding to one ormore of a commanded fuel flow to a combustion section, a variable bladepitch angle, a variable vane pitch angle, or a rotor plane angle. Incertain embodiments, the model-based controller, when included in thecomputing system with the sensor-based controller, is configured togenerate the output signal during substantially steady state engineoperating condition.

In still certain embodiments, the second set of operations executed bythe model-based controller includes receiving, via a sensor at theengine, an engine signal corresponding to one or more of a torque, a lowspool speed, a high spool speed, a rotor blade pitch angle, a vane pitchangle, a rotor plane position, or one or more actuator positionscorresponding to a variable geometry, or an acoustic sensor, andgenerating a thrust feedback signal based at least on the engine signal.

In certain embodiments, desired thrust and/or desired acoustic noiselevel is generated at least by adjusting the rotor blade pitch via arotor blade pitch change mechanism as a function of one or more of rotorassembly speed (e.g., N₁ speed), core engine speed (e.g., N₂ speed), andenvironmental conditions of the incoming air (including angle of attack,speed, temperature, pressure, humidity, etc.). In certain embodiments,desired thrust and/or acoustic noise level is generated at least byadjusting two or more parameter indicative of thrust output. Inparticular embodiments, the two or more parameters indicative of thrustoutput includes the rotor blade pitch at the rotor assembly and/or vanepitch angle at the vane assembly in aerodynamic relationship with therotor assembly, such as described herein. As such, thrust output may bealtered or generated by articulation of the vane assembly such asdescribed herein. Furthermore, operation of the engine may use fuel flowto the core engine and rotor blade pitch and/or vane pitch changes tocontrol rotor assembly (e.g., fan or propeller speed) in addition toengine pressure ratio, core sped, or rotor assembly torque. In someembodiments, methods for operating the engine include operating the coreengine at a substantially constant speed during one or more of groundoperation (e.g., ground idle, taxi, etc.), takeoff, climb, cruise,approach, landing, or thrust reverse. Operating the core engine at asubstantially constant speed may improve engine efficiency andperformance by allowing the core engine to operate substantially at orwithin an operating band (e.g., within 5%) of an aero design point(e.g., a design point for maximum performance in contrast to otheroperating conditions or speeds) for the core engine. Furthermore,operating the core engine at a substantially constant speed may contrastwith control systems or methods generally directed to operating anengine at a substantially constant fan or propeller speed whileadjusting gas generator or core speed.

In contrast to gas turbine engine configurations such as turbofans,embodiments of the engine provided herein allow for thrust control basedsubstantially on rotor blade pitch adjustment (e.g., via blade pitchchange mechanism 48), vane pitch adjustment (e.g., via vane pitchcontrol mechanism 148), or both. In certain embodiments, the core engine40 may operate at a substantially constant speed, such as to provideelectricity, air, or other services to an aircraft, such as for anenvironmental control system, a thermal management system, or poweringavionics or the aircraft generally. Operating the core engine (e.g.,core engine 40) at a substantially constant speed may allow embodimentsof the engine provided herein, as a propulsion system, to obviate aseparate auxiliary power unit (APU) for an aircraft. As such, aircraftweight and complexity may be reduced by allowing the propulsion unit toprovide power and services during ground operation that mayconventionally be provided by an APU.

In still certain embodiments, such as in regard to engine 10,substantially constant speed operation of the low speed spool 5054and/or the high speed spool 4046 may be allowed by rotating the rotorblade pitch 91 to reduce or increase thrust, including producing littleor no thrust from the rotor assembly when desired (e.g., groundoperations), substantially independent of speed of the core engine 40during operation of the engine 10. For example, the vane assembly 20 mayreduce or spoil thrust output from the rotor assembly such as describedin regard to FIGS. 20-21. Furthermore, or alternatively, thrust outputmay be adjusted by rotating the vane pitch 90 to reduce or increasethrust output, or to change thrust vector (e.g., reverse thrust, orprovide substantially axial thrust to correct for angle of attack). Assuch, in certain embodiments, methods and controllers provided hereinallow for thrust control via modulating or adjusting blade pitch angleand vane pitch angle while maintaining substantially constant speed ortorque of the core engine.

Embodiments of methods for control provided herein as may be stored orexecuted by one or more controllers 1600, 1610, 1700, 1800 may beutilized to desirably control rotor dynamics, such as vibrations, beatfrequencies, acoustics, etc. via trim controls or adjustments to bladepitch angle, vane pitch angle (e.g., altering loading at one or more ofthe rotor assembly and/or the vane assembly), or rotor plane angle. Instill various embodiments, methods and systems for control depicted anddescribed herein may include a supervisory controller configured as amodel-based controller. The supervisory controller may include an onlineoptimization program to improve fuel burn, perceived or measured noiselevels, combustor tones or dynamics, emissions output, or other controlor performance parameters, while maintaining desired thrust outputwithin operability limits.

Additionally, or alternatively, methods provided herein as may be storedor executed by one or more controllers 1600, 1610, 1700, 1800 may beutilized to desirably adjust or articulate blade pitch angle and/or vanepitch angle over a desired quantity of iterations, such as to determinea desired thrust output versus core engine speed, to remove or mitigateicing build up at a transient aircraft operating condition (e.g., duringtakeoff through icing conditions). In one embodiment, the controllerreceives an input signal indicative of an environmental parameter atwhich icing conditions may be present. The controller may generate anoutput signal corresponding to relatively rapid movements or changes inone or both of blade pitch angle or vane pitch angle to mitigateformation or build-up of ice at the rotor assembly or the vane assembly.The generated output signal may further be based at least on a torsionalmode shape of the blade 21 or the vane 31, such that the output signalcorresponds to a desired frequency (e.g., a resonance frequency). Themethod may include intermittent changes in pitch such as to remove icingbuild-up or mitigate ice build-up.

In addition to, or alternative to, steps of a method for operating anunducted rotor engine provided above, embodiments of the method oroperations provided herein may particularly provide for altering oradjusting thrust direction and magnitude, mitigating or eliminating beatfrequency, reducing undesired acoustics, reducing or removing ice ordebris build-up on a rotor assembly, and/or improve thrust match for anunducted rotor engine. Although embodiments of the method or operationsmay be applied with all or part of the embodiments of the engine 10provided herein, certain embodiments of the engine 10 may include acomputing system configured to execute one or more steps of the methodprovided herein. However, it should be appreciated that the embodimentsof the method, or portions thereof, may be executed with otherconfigurations of unducted rotor engine, gas turbine engine, orturbomachine.

In various embodiments, the method or operations includes operating acore engine and a rotor assembly to generate thrust output. In certainembodiments, such as the engine or portions thereof described herein,the core engine speed and the rotor assembly speed may be operatedsubstantially separately or at least partially independently of oneanother. For instance, blade pitch angles at one or more blades of therotor assembly may be altered to reduce or eliminate rotation of therotor assembly while the core engine is operating. As such, in certainembodiments, the method or operations includes determining a desiredthrust output versus speed of the core engine.

In still various embodiments, the method includes determining a desiredfirst blade pitch at the first blade of the rotor assembly anddetermining a desired second blade pitch at the second blade of therotor assembly. The method may include generating an output signal basedat least on the determined desired thrust output versus speed of thecore engine. In certain embodiments, such as described herein, thrustvector generated from the rotor assembly may be altered based at leaston operating the low speed spool (e.g., including the rotor assembly) ata substantially constant speed based at least on articulating one orboth of a first blade or a second blade of the variable pitch rotorassembly relative to a vane pitch angle of a vane assembly aft of therotor assembly.

The method includes adjusting or altering one or both of a first bladepitch at a first blade of the rotor assembly or a second blade pitch ata second blade of the rotor assembly based on a determined desiredthrust output versus speed of the core engine. In certain embodiments,the method includes articulating a first blade of a rotor assembly, suchdepicted and described herein, such as alter the first blade pitch whenarticulating the first blade. The method may further includearticulating a second blade of the rotor assembly different from thefirst blade. Articulating the second blade alters the second blade pitchof the second blade differently from the first blade pitch at the firstblade, such as depicted and described herein.

In various embodiments, the method includes receiving an input signalindicative of an environmental parameter within or surrounding thepropulsion system. The method or operations includes generating anoutput signal based on the environmental parameter. The output signalcorresponds to adjusting or articulating one or more of the first bladeor the second blade of the rotor assembly. In certain embodiments, theenvironmental parameter corresponds to one or more of a temperature,pressure, flow rate, density, or physical property of fluid entering theengine. In still certain embodiments, the environmental parametercorresponds to a perceived noise, ambient air temperature, ambient airpressure, or icing condition. The environmental parameter mayparticularly correspond to an operating altitude or attitude of anaircraft to which the engine is attached.

In still certain embodiments, generating the output signal correspondsto a desired frequency of articulation of one or more of the first bladeor the second blade of the rotor assembly. In one embodiment, thedesired frequency of articulation corresponds to a resonance frequencyof one or more of the first blade or the second blade of the rotorassembly. In another embodiment, the desired frequency of articulationis based at least on a torsional mode shape of the first blade or thesecond blade of the rotor assembly, such as described herein.

In yet another embodiment, articulating the first blade and the secondblade of the rotor assembly includes intermittently changing the firstblade pitch and the second blade pitch. For instance, intermittentchanging of the blade pitch may include actuating the first blade pitchand/or the second blade pitch each between a first angle (e.g.,theta_(1 first blade pitch), theta_(1 second blade pitch), etc.) and asecond angle (e.g., theta_(2 first blade pitch),theta_(2 second blade pitch)). In still certain embodiments,

In various embodiments, the method includes altering thrust vector basedat least on operating the core engine at a substantially constant speedand articulating one or both of the first blade or the second blade ofthe variable pitch rotor assembly. In one embodiment, the methodincludes altering thrust vector based at least on operating the lowspeed spool at a substantially constant speed based at least onarticulating one or both of the first blade or the second blade of thevariable pitch rotor assembly relative to a vane pitch angle of the vaneassembly aft of the rotor assembly. In another embodiment, the methodincludes altering thrust vector based at least on operating the highspeed spool at a substantially constant speed based at least onarticulating one or both of the first blade or the second blade of thevariable pitch rotor assembly relative to a vane pitch angle of the vaneassembly.

Various embodiments of the method may be executed with a controller(e.g., computing system 210 in FIG. 1) in which all or part of themethod is stored as instructions and/or executed as operations at one ormore embodiments of an engine, such as the engine 10 depicted anddescribed herein. In certain embodiments, the method is performed at anengine including a single unducted rotor assembly positioned forward ofa vane assembly. In still certain embodiments, the method is executed atan engine including an unducted rotor assembly positioned in aerodynamicrelationship with a variable pitch vane assembly.

Still various embodiments of the method may be executed with an engineincluding a variable pitch rotor assembly including a single stage of aplurality of blades coupled to a disk in which the plurality of bladesincludes a first blade configured to articulate a first blade pitchseparately from a second blade configured to articulate a second bladepitch. A fixed-pitch or variable-pitch vane assembly may be positionedforward or aft of the variable pitch rotor assembly. The engine includesa core engine including a high speed spool and a low speed spool isoperably coupled to the rotor assembly.

Referring now to FIGS. 30-31, an aircraft with symmetric open rotorengine configurations (e.g., left wing engines 10B and right wingengines 10A, or left-side fuselage engines 10B and right-side fuselageengines 10A, etc.) may be susceptible to undesired acoustic noise basedon acoustic beat interferences due to differences in rotor assemblyfrequencies between the plurality of engines. It should be appreciatedthat beat interference or beat frequency is an interference pattern fromtwo or more engines operating at different frequencies, perceived as aperiodic variation in volume whose rate is the difference between thetwo or more frequencies from the respective engines.

The controllers depicted and described in regard to FIGS. 23-29 may beconfigured substantially similarly as shown and described in regard toFIGS. 30-31. In the embodiment depicted in FIG. 28, a first enginedepicted as Engine 1 or engine 10A represents one or more right-wing orright-side fuselage engines, and a second engine depicted as Engine 2 orengine 10B represents one or more left-wing or left-side fuselageengines. The engine controllers from Engine 1 and Engine 2 are coupledtogether in analog or digital communication. Cross coupling of the twoor more engine controllers allows for communication between therespective engine controllers (e.g., controller 210 for each respectiveengine, such as depicted in FIG. 1) of an engine operating parameter,such as one or more of the rotor assembly rotational speed, rotor bladepitch angle, vane pitch angle, or rotor plane (e.g., via an Angle ofAttack control mechanism). In alternate embodiments, this cross-couplingor communication between engines may be via the aircraft flight controland communication bus.

In the embodiment depicted in FIG. 29, two or more of the enginecontrols from Engine 1 and Engine 2 are coupled together in analog ordigital communication via a master controller 1900. The mastercontroller 1900 is configured to receive inputs from Engine 1 and Engine2 and determine whether to adjust an engine operating parameter atEngine 1, at Engine 2, or both, to synchrophase the plurality ofengines. In certain embodiments, the master controller 1900 is adesignated engine controller from either engine 1 or engine 2 (e.g., anengine controller, such as a Full Authority Digital Engine Controller,or FADEC). In other embodiments, the master controller is an aircraftcontroller, such as via aircraft avionics. In yet another embodiment,the master controller is, or is a portion of, a distributed networkconfigured to receive and transmit signals from and to the two or moreengines.

In various embodiments, the designated engine controller is altered orvarying between Engine 1 and Engine 2. In such an embodiment, thedesignated engine controller (e.g., the master controller 1900) may bevaried based on a relative engine performance of Engine 1 and Engine 2.In one embodiment, the designated engine controller is based at least ona better-performing engine, in which the better-performing engine isbased on one or more of a health parameter, an engine operatingparameter, an engine cycle count, an exhaust gas temperature, specificfuel consumption, time-on-wing, or other desired parameter establishingone engine as determinative of an engine operating condition to whichanother engine will be adjusted to.

In another embodiment, the designated engine controller is based atleast on a least-performing engine, such as to define a lowest commondenominator of engine performance between Engine 1 and Engine 2. In suchan embodiment, the better-performing engine (i.e., not the leastperforming engine) may be de-tuned, de-rated, or otherwise adjusted tosubstantially match the engine operating condition of theleast-performing engine.

Referring back to FIGS. 30-31, the controller and method determinesdifferences in the engine operating parameter. In certain embodiments,the controller and method further compare the differences in engineoperating parameter to one or more of a desired acoustic noise level, adesired thrust output, a health parameter, or a performance parameter(e.g., specific fuel consumption). The controller and method determineone or more adjustments, trims, or other changes to the engine operatingparameter based at least on changing rotor assembly (e.g., rotorassembly 20) at one or more engines to match the plurality of engines atthe aircraft to mitigate undesired noise. The controller and method mayfurther determine one or more adjustments to the engine operatingparameter to mitigate or eliminate undesired noise while avoidingasymmetric thrust or power conditions.

The controller and method further adjust the engine operating parameterbased on the determined adjustment. In certain embodiments, the methodincludes decreasing the performance parameter at a first engine tosubstantially match the performance parameter at a second engine andfurther synchrophase rotor speeds and outputs to reduce or eliminatebeat interferences. In another embodiment, a first engine including abetter health parameter may be de-tuned to match a second engineincluding a relatively worse health parameter to reduce or eliminatebeat interferences. In still another embodiment, a first engineincluding certain desired levels of health parameter or performanceparameter may increase or decrease rotor assembly speed, pitch angle,etc. to mitigate or attenuate beat interferences while reducing thehealth parameter or performance parameter (e.g., reducing within stillacceptable or desired limits).

It should be appreciated that embodiments of the method, controller,engine, or aircraft provided herein may include cross coupling two ormore engines to improve overall aircraft and system performance andmitigate or eliminate undesired noise, such as beat frequency, whilepreserving desired thrust output. As such, although certain embodimentsmay include de-rating or de-tuning an engine to match another engine, itshould be appreciated that the two or more engines may definesubstantially similar performance characteristics. Embodiments of themethod and system provided herein may receive engine parameters from therespective engines and adjust one or more of the core speed (e.g., fuelflow, engine loading via an electric machine or variable vanes, valves,orifices, etc., high spool speed, etc.) or rotor blade pitch angle. Thesystem may additionally adjust vane pitch angle and/or rotor plane angleto improve acoustic noise levels based on other noise sources (e.g.,those not based on beat frequency), or to control thrust output level,or to control thrust output vector (e.g., providing a more axial thrust,such as to allow for lower core engine speeds and fuel consumption whilemaintaining or increasing thrust output at the rotor assembly and vaneassembly).

Various embodiments of the engine 10 depicted and described hereinprovide novel improvements over known propulsion systems. Embodiments ofthe engine 10 include, but are not limited to, one or more ranges ofratios of blades to vanes, length to maximum diameter, vane spacing ororientation (i.e., vane pitch angle) relative to one or more blades orblade pitch angle, or combinations thereof. It should be appreciatedthat, to the extent one or more structures or ranges may overlap one ormore of those known in the art, certain structures with certain turbomachine arrangements may be generally undesired to combine with otherstructures of other turbo machine arrangements. For instance, turbofanconfigurations generally include certain quantities of vanes to providestructural support for a casing surrounding a rotor assembly, withoutproviding any teaching or motivation in regard to thrust output andnoise abatement particular to open rotor engines. In another instance,turboprop or turboshaft configurations generally exclude vane assembliessince the added structure may increase weight without providing otherbenefits for turboprop or turbofan applications.

In still another instance, certain ranges of blades to vanes describedherein provide unexpected benefits not previously known in the art, orfurthermore, not previously known in the art for single stage unductedrotor assemblies. In still yet another instance, certain ranges ofblades to vanes with certain ranges of length to maximum diameter of theengine provide unexpected benefits not previously known in the art, orfurthermore, not previously known in the art for single stage unductedrotor assemblies. In still particular embodiments, certain ranges,differences, or sums of blades and vanes provided herein provideunexpected benefits not previously known in the art, such as reducedinteraction noise between the blade assembly 20 and the vane assembly30.

Still further, certain embodiments of the engine 10 provided herein mayallow for normal subsonic aircraft cruise altitude operation at or aboveMach 0.5, or above Mach 0.75, based at least on ranges or quantities ofblades to vanes and/or ranges of blades to vanes and length to maximumdiameter, and/or in combination with other structures provided herein.In certain embodiments, the engine 10 allows for normal aircraftoperation between Mach 0.55 and Mach 0.85, or between Mach 0.75 to Mach0.85 at cruise altitude. In certain embodiments, the engine 10 allowsfor rotor blade tip speeds at or less than 750 feet per second (fps). Instill certain embodiments, the core engine 40 and rotor assembly 20 aretogether configured to produce a threshold power loading is 25horsepower per ft² or greater at cruise altitude. In particularembodiments of the engine 10, structures and ranges provided hereingenerate power loading between 25 horsepower/ft² and 100 horsepower/ft²at cruise altitude. Still particular embodiments may provide suchbenefits with reduced interaction noise between the blade assembly 20and the vane assembly 30 and/or decreased overall noise generated by theblade assembly 20 and the vane assembly 30. Additionally, it should beappreciated that ranges of power loading and/or rotor blade tip speedmay correspond to certain structures, core sizes, thrust outputs, etc.,or other structures at the core engine 40 and the rotor assembly 20.However, as previously stated, to the extent one or more structuresprovided herein may be known in the art, it should be appreciated thatthe present disclosure may include combinations of structures notpreviously known to combine, at least for reasons based in part onconflicting benefits versus losses, desired modes of operation, or otherforms of teaching away in the art.

It should furthermore be appreciated that certain unexpected benefits ofvarious embodiments of the engine 10 provided herein may provideparticular improvements to propulsion systems in regard to thrust outputand acoustic levels. For instance, quantities of blades greater thanthose of one or more ranges provided herein may produce noise levelsthat may disable use of an open rotor engine in certain applications(e.g., commercial aircraft, regulated noise environments, etc.). Inanother instance, quantities of blades less than those ranges providedherein may produce insufficient thrust output, such as to render an openrotor engine non-operable in certain aircraft applications. In yetanother instance, quantities of vanes less than those of one or moreranges provided herein may fail to sufficiently produce thrust and abatenoise, such as to disable use of an open rotor engine in certainapplications. In still another instance, quantities of vanes greaterthan those of ranges provided herein may result in increased weight thatadversely affects thrust output and noise abatement.

It should be appreciated that embodiments of the engine 10 including oneor more ranges of ratios, differences, sums, or discrete quantities ofblades 21 to vanes 31 depicted and described herein may provideadvantageous improvements over turbofan or turboprop gas turbine engineconfigurations. In one instance, embodiments of the engine 10 providedherein allow for thrust ranges similar to or greater than turbofanengines with larger quantities of blades or vanes, while furtherobviating structures such as fan cases or nacelles. In another instance,embodiments of the engine 10 provided herein allow for thrust rangessimilar to or greater than turboprop engines with similar quantities ofblades, while further providing reduced noise or acoustic levels such asprovided herein. In still another instance, embodiments of the engine 10provided herein allow for thrust ranges and attenuated acoustic levelssuch as provided herein while reducing weight, complexity, or issuesassociated with fan cases, nacelles, variable nozzles, orthrust-reverser assemblies at a turbofan nacelle.

This written description uses examples to disclose the invention,including the best mode, and also to enable any person skilled in theart to practice the invention, including making and using any devices orsystems and performing any incorporated methods. The scope of theinvention(s) described herein is defined by one or more of the claims,including combinations of two or more claims or clauses (as set forthbelow) and may include other examples that occur to those skilled in theart. For example, aspects of the invention(s) are provided by thesubject matter of the following clauses, which are intended to cover allsuitable combinations unless dictated otherwise based on logic or thecontext of the clauses and/or associated figures and description:

1. A propulsion system defining an engine centerline, the propulsionsystem comprising a rotor assembly comprising a plurality of bladesextended radially relative to the engine centerline axis; and a vaneassembly positioned in aerodynamic relationship with the rotor assembly,wherein the vane assembly comprises a plurality of vanes extendedradially relative to the engine centerline axis, and wherein thepropulsion system comprises a ratio of a quantity of blades to aquantity of vanes between 2:5 and 2:1.

2. The propulsion system of any one or more clauses herein, wherein thequantity of blades is 20 or fewer.

3. The propulsion system of any one or more clauses herein, wherein thequantity of blades is between 16 and 11.

4. The propulsion system of any one or more clauses herein, wherein adifference between the quantity of vanes and the quantity of blades isbetween 2 and −2.

5. The propulsion system of any one or more clauses herein, wherein adifference between the quantity of vanes and the quantity of blades isbetween 2 and −2, and wherein the quantity of blades is between 16 and11.

6. The propulsion system of any one or more clauses herein, wherein theratio of the quantity of blades to the quantity of vanes between 0.5 and1.5.

7. The propulsion system of any one or more clauses herein, wherein asum of blades and vanes is 30 or fewer, and wherein the sum of bladesand vanes is 20 or greater.

8. The propulsion system of any one or more clauses herein, wherein therotor assembly is unducted.

9. The propulsion system of any one or more clauses herein, wherein thevane assembly is positioned aft of the rotor assembly.

10. The propulsion system of any one or more clauses herein, wherein thevane assembly is unducted.

11. The propulsion system of any one or more clauses herein, thepropulsion system comprising a core engine encased in a nacelle, whereinthe nacelle defines a maximum diameter, and wherein the vane assembly isextended from the nacelle.

12. The propulsion system of any one or more clauses herein, wherein therotor assembly comprises a hub from which the plurality of blades isextended, and wherein the propulsion system comprises a length extendedfrom a forward end of the hub to an aft end of the nacelle, and whereina ratio of length to maximum diameter is at least 2.

13. The propulsion system of any one or more clauses herein, wherein theratio of length to maximum diameter is at least 2.5.

14. The propulsion system of any one or more clauses herein, wherein thecore engine and the rotor assembly are together configured to generate apower loading of 25 horsepower per square foot or greater at cruisealtitude.

15. The propulsion system of any one or more clauses herein, wherein therotor assembly is configured to rotate at a blade tip speed of up to 750feet per second.

16. A propulsion system defining an engine centerline axis, thepropulsion system comprising an unducted single rotor assemblycomprising a plurality of blades extended radially relative to theengine centerline axis; and a vane assembly positioned aft of theunducted rotor assembly, wherein the vane assembly comprises a pluralityof vanes extended radially relative to the engine centerline axis, andwherein the propulsion system comprises a difference between a quantityof vanes and a quantity of blades is between 2 and −2.

17. The propulsion system of any one or more clauses herein, wherein therotor assembly comprises a blade pitch change mechanism configured tocontrol blade pitch at one or more of the plurality of blades relativeto vane pitch at one or more of the plurality of vanes.

18. The propulsion system of any one or more clauses herein, wherein thevane assembly comprises a vane pitch change mechanism configured tocontrol vane pitch at one or more of the plurality of vanes relative toblade pitch at one or more of the plurality of blades.

19. The propulsion system of any one or more clauses herein, thepropulsion system comprising a core engine encased in a nacelle, whereinthe nacelle defines a maximum diameter, and wherein the vane assembly isextended from the nacelle; and wherein the rotor assembly comprises ahub from which the plurality of blades is extended, and wherein thepropulsion system comprises a length extended from a forward end of thehub to an aft end of the nacelle, and wherein a ratio of length tomaximum diameter is at least 2.

20. The propulsion system of any one or more clauses herein, wherein asum of blades and vanes is 30 or fewer, and wherein the sum of bladesand vanes is 20 or greater.

21. The propulsion system in accordance with one or more clauses of thischapter, wherein the propulsion system generates a power loading at therotor assembly of 50 horsepower per ft² or less at cruise altitude.

22. A propulsion system, the propulsion system comprising a core engineencased in a nacelle, wherein the nacelle defines a diameter, a rotorassembly comprising a plurality of blades and a hub, a vane assemblyextended from the nacelle of the core engine, the vane assemblypositioned aft of the rotor assembly, the propulsion system defines alength extended from the hub of the rotor assembly to an aft end of thenacelle, and wherein a ratio of length to diameter is at least 2, or atleast 2.5

23. The propulsion system in accordance with one or more clauses of thischapter, wherein the core engine comprises an axisymmetric inlet.

24. The propulsion system in accordance with one or more clauses of thischapter comprising the system for reducing noise generation for a singleunducted rotor engine in accordance with one or more clauses of thischapter.

25. A propulsion system, the propulsion system comprising a variablepitch rotor assembly comprising a plurality of blades coupled to a disk,wherein the plurality of blades extends radially from the disk, andwherein the plurality of blades is positioned along a rotor plane, therotor plane extended orthogonal to a longitudinal centerline axis of therotor assembly; and a scheduling ring rotatable relative to the disk andhaving a plurality of slots, and a plurality of linkage arms, eachlinkage arm operatively coupled to one of the plurality of fan bladesand to one of the plurality of slots, wherein each of the plurality offan blades rotates according to a blade pitch schedule defined by theslot to which it is operatively coupled, and wherein at least two of theplurality of slots define different blade pitch schedules.

26. The propulsion system in accordance with one or more clauses of thischapter, wherein at least two of the plurality of slots define differentblade pitch schedules.

27. The propulsion system in accordance with one or more clauses of thischapter, wherein each of the plurality of linkage arms has a first endfixedly connected to one of the plurality of blades and a second endslidably connected to one of the plurality of slots.

28. The propulsion system in accordance with one or more clauses of thischapter, wherein the plurality of blades comprises a first set of bladesand a second set of blades, and wherein the first set of blades isoperably coupled to a first scheduling slot defining a first bladeschedule, and wherein the second set of blades is operably coupled to asecond scheduling slot defining a second blade schedule different fromthe first blade schedule.

29. The propulsion system in accordance with one or more clauses of thischapter, wherein the first scheduling slot and the second schedulingslot are in adjacent alternating arrangement.

30. The propulsion system in accordance with one or more clauses of thischapter, wherein the rotor assembly is a single unducted rotor assemblyconfigured to provide substantially axial thrust.

31. The propulsion system in accordance with one or more clauses of thischapter, wherein the rotor assembly comprises between eight and twentyblades rotatably coupled to the disk.

32. The propulsion system in accordance with one or more clauses of thischapter, wherein the rotor assembly comprises twelve blades.

33. The propulsion system in accordance with one or more clauses of thischapter, comprising a core engine configured to produce combustion gasesfor driving a turbine section, wherein the variable pitch rotor assemblyis configured to provide changes in thrust vector without changes inspeed at the core engine.

34. The propulsion system in accordance with one or more clauses of thischapter comprising the system for reducing noise generation for a singleunducted rotor engine in accordance with one or more clauses of thischapter.

75. A method for thrust reverse for a single unducted rotor engine, themethod comprising generating a forward velocity component forward of arotor assembly at least by closing a blade pitch angle at a blade at therotor assembly, and generating a positive tangential velocity componentat the rotor assembly, wherein the positive tangential velocitycomponent is opposite of a negative tangential velocity component from avane assembly aft of the rotor assembly.

76. The method in accordance with one or more clauses of this chapter,comprising loading the rotor assembly at least by closing a vane pitchangle at a vane at the vane assembly.

77. The method in accordance with one or more clauses of this chapter,wherein loading the rotor assembly comprises rotating the vane at thevane assembly co-directional to a direction of rotation of the blade atthe blade assembly.

78. The method in accordance with one or more clauses of this chapter,wherein loading the rotor assembly comprises reducing the negativetangential velocity component at the vane assembly at least by closingthe vane pitch angle.

79. A method for adjusting thrust vector for a single unducted rotorengine, the method comprising one or more steps of any preceding clause.

80. A computer-implemented method for operating a single unducted rotorengine, the computer-implemented method comprising the method of anypreceding clause.

81. A computing system comprising one or more processors and one or morememory devices, wherein the one or more memory devices is configured tostore instructions that, when executed by the one or more processors,performs operations, the operations comprising any of the steps of themethod of any preceding clause.

82. A propulsion system in accordance with one or more clauses of thischapter, the propulsion system comprising the computing system of anypreceding clause.

83. A propulsion system in accordance with one or more clauses of thischapter, the propulsion system configured to execute the steps of themethod in accordance with one or more clauses of this chapter.

84. A computing system comprising one or more processors and one or morememory devices, wherein the one or more memory devices is configured tostore instructions that, when executed by the one or more processors,performs operations, the operations comprising obtaining a throttleinput corresponding to one or more of a desired air speed of anaircraft, thrust output, or pressure ratio; generating, via an enginecycle model, a commanded rotor blade pitch angle; obtaining, via asensor, a measured rotor blade pitch; and generating, via a controller,a rotor blade pitch signal corresponding to a commanded adjustment tothe commanded rotor blade pitch angle based at least on the measuredrotor blade pitch.

85. The computing system in accordance with one or more clauses of thischapter, comprising generating a corrected low spool parameter signalbased at least on a commanded low spool parameter and a measured lowspool parameter; and generating an engine control signal correspondingto a fuel flow at a combustion section.

86. The computing system in accordance with one or more clauses of thischapter, wherein generating the rotor blade pitch signal is a firstcontrol loop and generating the engine control signal is a secondcontrol loop.

87. The computing system in accordance with one or more clauses of thischapter, wherein generating the rotor blade pitch signal is independentof generating the engine control signal when the combustion section isat a substantially steady state aircraft operating condition

88. A computing system comprising one or more processors and one or morememory devices, wherein the one or more memory devices is configured tostore instructions that, when executed by the one or more processors,performs operations, the operations comprising obtaining a desiredthrust output via a throttle input; determining, at least via a powermanagement logic block, a commanded thrust output signal; receiving, ata controller, the commanded thrust output signal; and generating, viathe controller, an output signal corresponding to one or more of acommanded fuel flow, a rotor blade pitch angle, a vane pitch angle, or arotor plane angle.

89. The computing system in accordance with one or more clauses of thischapter, comprising generating, via the engine, an actual thrust outputestimate based at least on the output signal.

90. The computing system in accordance with one or more clauses of thischapter, comprising receiving, via a sensor at the engine, an enginesignal corresponding to one or more of a torque measurement, a low spoolspeed measurement, a high spool speed measurement, a rotor blade pitchangle measurement, a vane pitch angle measurement, a rotor planeposition measurement, or one or more actuator positions corresponding toa variable geometry, or an acoustic sensor.

91. The computing system in accordance with one or more clauses of thischapter, comprising generating a thrust feedback signal based at leaston the engine signal.

92. The computing system in accordance with one or more clauses of thischapter, wherein the thrust curve defines engine output characteristicsas a function of one or more environmental conditions or control devicesat the engine.

93. A computing system comprising one or more processors and one or morememory devices, wherein the one or more memory devices is configured tostore instructions that, when executed by the one or more processors,performs operations, the sensor-based controller configured to executeinstructions that perform operations comprising obtaining a throttleinput corresponding to one or more of a desired air speed of anaircraft, thrust output, or pressure ratio; generating, via an enginecycle model, a commanded rotor blade pitch angle; obtaining, via asensor, a measured rotor blade pitch; and generating, via a controller,a rotor blade pitch signal corresponding to a commanded adjustment tothe commanded rotor blade pitch angle based at least on the measuredrotor blade pitch, and wherein the computing system comprises amodel-based controller configured to execute operations comprisingobtaining a desired thrust output via a throttle input; determining, atleast via an engine cycle model comprising a thrust curve, a commandedthrust output signal; receiving, at a controller, the commanded thrustoutput signal; and generating, via the controller, an output signalcorresponding to one or more of a commanded fuel flow, a rotor bladepitch angle, a vane pitch angle, or a rotor plane angle.

94. The computing system in accordance with one or more clauses of thischapter, wherein the model-based controller is configured to adjustengine output thrust within a 7% margin.

95. The computing system in accordance with one or more clauses of thischapter, wherein the model-based controller is configured to adjustengine output thrust within a 5% margin.

96. The computing system in accordance with one or more clauses of thischapter, wherein the sensor-based controller is configured to generatethe rotor pitch signal during transient changes in engine operatingcondition.

97. The computing system in accordance with one or more clauses of thischapter, wherein transient changes in engine operating conditioncomprises conditions to and between two or more of ignition, idle,takeoff, climb, cruise, descent, approach, or thrust reverse.

98. The computing system in accordance with one or more clauses of thischapter, wherein the model based controller is configured to generatethe output signal during substantially steady state engine operatingcondition.

99. The computing system in accordance with one or more clauses of thischapter, the computing system configured to execute operations, theoperations comprising generating a control signal corresponding to acommanded engine pressure ratio, a commanded core engine pressure ratio,a commanded rotor blade pitch angle, a commanded fuel flow, a commandedrotor plane angle, a commanded vane pitch angle, or combinationsthereof.

100. The computing system in accordance with one or more clauses herein,the computing system comprising a first engine controller correspondingto a first single unducted rotor engine, the first engine controllercomprising the sensor-based controller and the model-based controller; asecond engine controller corresponding to a second single unducted rotorengine, the second engine controller comprising the sensor-basedcontroller and the model-based controller, wherein the first enginecontroller and the second engine controller are in cross coupledcommunication to one another, and wherein the operations comprisedetermining a master controller of the first engine controller and thesecond engine controller, and determining whether to adjust an engineoperating parameter at one or more of the first single unducted rotorengine or the second single unducted rotor engine, wherein determiningwhether to adjust the engine operating parameter corresponds to a sensedbeat frequency between the first and second single unducted rotorengines.

101. The computing system of any clause herein, the operationscomprising adjusting one or more of an engine operating parametercomprising rotor blade pitch angle, a vane pitch angle, or a rotor planeangle at one or more of the first single unducted rotor engine or thesecond single unducted rotor engine.

102. The computing system of any clause herein, wherein adjusting one ormore of the engine operating parameter comprises maintaining asubstantially constant speed of a core engine at the single unductedrotor engines.

103. The computing system of any clause herein, wherein determining themaster controller comprises determining a better-performing engine basedon one or more of a health parameter, an engine operating parameter, anengine cycle count, an exhaust gas temperature, specific fuelconsumption, or time-on-wing.

104. The computing system of any clause herein, the operationscomprising de-tuning the better-performing single unducted rotor enginebased on the other engine operating condition.

105. The computing system of any clause herein, the operationscomprising generating, via the sensor-based controller at the firstengine controller or the second engine controller, a corrected low spoolparameter signal based at least on a commanded low spool parameter and ameasured low spool parameter; generating, via the sensor-basedcontroller at the first engine controller or the second enginecontroller, an engine control signal corresponding to a fuel flow at acombustion section; generating, via the engine, an actual thrust outputbased at least on the output signal; and receiving, via a sensor at theengine, an engine signal corresponding to one or more of a torquemeasurement, a low spool speed measurement, a high spool speedmeasurement, a rotor blade pitch angle measurement, a vane pitch anglemeasurement, a rotor plane position measurement, or one or more actuatorpositions corresponding to a variable geometry, or an acoustic sensor.

106. An aircraft comprising a computing system in accordance with one ormore clauses of this chapter.

107. An aircraft comprising a propulsion system in accordance with oneor more clauses of this chapter.

108. An aircraft configured to execute the steps of a method inaccordance with one or more clauses of this chapter.

109. A propulsion system, the propulsion system comprising a variablepitch rotor assembly comprising a plurality of blades coupled to a disk,wherein the plurality of blades comprises a first blade configured toarticulate a first blade pitch separately from a second blade configuredto articulate a second blade pitch; a vane assembly positioned inaerodynamic relationship with the variable pitch rotor assembly; a coreengine comprising a high speed spool and a low speed spool, wherein thelow speed spool is operably coupled to the rotor assembly; and one ormore controllers configured to execute operations, the operationscomprising articulating the first blade of the rotor assembly, whereinarticulating the first blade alters the first blade pitch; andarticulating the second blade of the rotor assembly, whereinarticulating the second blade alters the second blade pitch.

110. The propulsion system of any clause herein, the operationscomprising receiving an input signal indicative of an environmentalparameter within or surrounding the propulsion system; and generating anoutput signal based on the environmental parameter, wherein the outputsignal corresponds to articulation of one or more of the first blade orthe second blade of the rotor assembly.

111. The propulsion system of any clause herein, wherein generating theoutput signal corresponds to a desired frequency of articulation of oneor more of the first blade or the second blade of the rotor assembly.

112. The propulsion system of any clause herein, wherein the desiredfrequency of articulation corresponds to a resonance frequency of one ormore of the first blade or the second blade of the rotor assembly.

113. The propulsion system of any clause herein, wherein the desiredfrequency of articulation is based at least on a torsional mode shape ofthe first blade or the second blade of the rotor assembly.

114. The propulsion system of any clause herein, wherein articulatingthe first blade and the second blade of the rotor assembly comprisesintermittently changing the first blade pitch and the second bladepitch.

115. The propulsion system of any clause herein, the operationscomprising altering thrust vector based at least on operating the coreengine at a substantially constant speed and articulating one or both ofthe first blade or the second blade of the variable pitch rotorassembly.

116. The propulsion system of any clause herein, the operationscomprising altering thrust vector based at least on operating the lowspeed spool at a substantially constant speed based at least onarticulating one or both of the first blade or the second blade of thevariable pitch rotor assembly relative to a vane pitch angle of the vaneassembly aft of the rotor assembly.

117. The propulsion system of any clause herein, wherein the variablepitch rotor assembly is a single unducted rotor assembly positionedforward of the vane assembly.

118. The propulsion system of any clause herein, the operationscomprising operating the core engine and the rotor assembly to generatethrust output; determining a desired thrust output versus speed of thecore engine; and generating an output signal based at least on thedetermined desired thrust output versus speed of the core engine.

119. The propulsion system of any clause herein, the operationscomprising altering thrust vector based at least on operating the lowspeed spool at a substantially constant speed based at least onarticulating one or both of the first blade or the second blade of thevariable pitch rotor assembly relative to a vane pitch angle of the vaneassembly aft of the rotor assembly.

120. A propulsion system comprising a variable pitch rotor assemblycomprising a single stage of a plurality of blades coupled to a disk,wherein the plurality of blades comprises a first blade configured toarticulate a first blade pitch separately from a second blade configuredto articulate a second blade pitch; a vane assembly positioned aft ofthe variable pitch rotor assembly; a core engine comprising a high speedspool and a low speed spool, wherein the low speed spool is operablycoupled to the rotor assembly; and a controller configured to executeoperations, the operations comprising determining a desired thrustoutput versus speed of the core engine; determining a desired firstblade pitch at the first blade; determining a desired second blade pitchat the second blade; and adjusting one or both of the first blade pitchor the second blade pitch based on the determined desired thrust outputversus speed of the core engine.

121. The propulsion system of any clause herein, the operationscomprising receiving an input signal indicative of an environmentalparameter; and generating an output signal based on the environmentalparameter, wherein the output signal corresponds to adjusting of one ormore of the first blade pitch or the second blade pitch.

122. The propulsion system of any clause herein, wherein theenvironmental parameter comprises one or more of a perceived noise,ambient air temperature, ambient air pressure, or icing condition.

123. The propulsion system of any clause herein, wherein generating theoutput signal corresponds to a desired frequency of adjusting of one ormore of the first blade pitch or the second blade pitch.

124. The propulsion system of any clause herein, wherein the desiredfrequency of adjusting corresponds to a resonance frequency of one ormore of the first blade or the second blade of the rotor assembly.

125. The propulsion system of any clause herein, wherein the desiredfrequency of adjusting is based at least on a torsional mode shape ofthe first blade or the second blade of the rotor assembly.

126. The propulsion system of any clause herein, wherein adjusting oneor both of the first blade pitch or the second blade pitch comprisesintermittently adjusting the first blade pitch and the second bladepitch between a respective first angle and second angle.

127. The propulsion system of any clause herein, the operationscomprising altering thrust vector based at least on adjusting the firstblade to the desired first blade pitch at the first blade or adjustingthe second blade to the desired second blade pitch different from thedesired first blade pitch.

128. The propulsion system of any clause herein, the operationscomprising operating the core engine at a substantially constant speedwhen altering thrust vector.

129. A propulsion system defining an engine centerline, the propulsionsystem comprising a rotor assembly configured to rotate relative to theengine centerline axis, the rotor assembly comprising a plurality ofblades, each blade of the plurality of blades configured to rotate alonga respective blade pitch angle axis; and a vane assembly positioned inaerodynamic relationship with the rotor assembly, the vane assemblycomprising a plurality of vanes, each vane of the plurality of vanesconfigured to rotate along a respective vane pitch angle axis acontroller configured to execute operations, the operations comprisingmoving each blade of the plurality of blades to a reverse thrustposition about its respective blade pitch axis, wherein a leading edgeof each blade is located aft of a trailing edge of the respective bladeat a radial span location when in the reverse thrust position; andadjusting each vane of the plurality of vanes about its respective vanepitch axis when the plurality of blades is in the reverse thrustposition to modify an amount of reverse thrust generated by thepropulsion system.

130. The propulsion system of any clause herein, wherein the rotorassembly is unducted.

131. The propulsion system of any clause herein, wherein the one or moreblades is configured to generate forward flow over a first portion of ablade span, and wherein the one or more blades is configured to generatereverse flow over a second portion of the blade span.

132. The propulsion system of any clause herein, wherein the one or moreblades is configured to generate forward flow below 50% of a blade span,and wherein the one or more blades is configured to generate reverseflow at or above 50% of the blade span.

133. The propulsion system of any clause herein, wherein adjusting eachvane of the plurality of vanes about its respective vane pitch axis whenthe plurality of blades are in the reverse thrust position comprisesrotating one or more vanes along the vane pitch axis up to 15 degreesopen or up to 15 degrees closed from a design point.

134. The propulsion system of any clause herein, wherein adjusting eachvane of the plurality of vanes about its respective vane pitch axis whenthe plurality of blades are in the reverse thrust position comprisesrotating one or more vanes along the vane pitch axis up to 10 degreesopen or up to 10 degrees closed from a design point.

135. The propulsion system of any clause herein, wherein adjusting eachvane of the plurality of vanes about its respective vane pitch axis whenthe plurality of blades are in the reverse thrust position comprisesrotating one or more vanes along the vane pitch axis up to 5 degreesopen or up to 5 degrees closed from a design point.

136. The propulsion system of any clause herein, wherein adjusting eachvane of the plurality of vanes about its respective vane pitch axis whenthe plurality of blades is in the reverse thrust position comprisesclosing the vanes to increase the amount of reverse thrust generated bythe propulsion system.

137. The propulsion system of any clause herein, wherein adjusting eachvane of the plurality of vanes about its respective vane pitch axis whenthe plurality of blades is in the reverse thrust position comprisesopening the vanes to decrease the amount of reverse thrust generated bythe propulsion system.

138. The propulsion system of any clause herein, wherein the rotorassembly is ducted.

139. The propulsion system of any clause herein, wherein the vaneassembly is positioned aft of the rotor assembly when the blade pitchangle at one or more blades of the rotor assembly is closed.

140. A method for generating reverse thrust for a single stage unductedrotor engine with a vane assembly positioned in aerodynamicrelationship, the method comprising adjusting a blade pitch angle at oneor more blades of the rotor assembly to position a blade leading edgeaft of a blade trailing edge at a radial span location; and adjustingloading at the rotor assembly based on changing a vane pitch angle ofone or more vanes of the vane assembly.

141. The method of any clause herein, wherein adjusting loading at therotor assembly based on changing the vane pitch angle of one or morevanes of the vane assembly comprises closing the vanes to increase theamount of reverse thrust generated by the engine.

142. The method of any clause herein, wherein adjusting loading at therotor assembly based on changing the vane pitch angle of one or morevanes of the vane assembly comprises opening the vanes to decrease theamount of reverse thrust generated by the propulsion system.

143. The method of any clause herein, wherein adjusting loading at therotor assembly based on changing the vane pitch angle of one or morevanes of the vane assembly comprises rotating one or more vanes alongthe vane pitch axis up to 15 degrees open or up to 15 degrees closedfrom a design point.

144. The method of any clause herein, wherein adjusting loading at therotor assembly based on changing the vane pitch angle of one or morevanes of the vane assembly comprises rotating one or more vanes alongthe vane pitch axis up to 10 degrees open or up to 10 degrees closedfrom a design point.

145. The method of any clause herein, wherein adjusting loading at therotor assembly based on changing the vane pitch angle of one or morevanes of the vane assembly comprises rotating one or more vanes alongthe vane pitch axis up to 5 degrees open or up to 5 degrees closed froma design point.

146. The method of any clause herein, wherein adjusting the blade pitchangle at one or more blades of the rotor assembly to position the bladeleading edge aft of the blade trailing edge at the radial span locationcomprises generating forward flow below 50% of a blade span; andgenerating reverse flow at or above 50% of the blade span.

147. A computing system, the computing system configured to storeinstructions that, when executed by the one or more processors, performsoperations, the operations comprising commanding an adjustment of ablade pitch angle at one or more blades of a rotor assembly of anaeronautical engine to position a blade leading edge aft of a bladetrailing edge at a radial span location; and commanding an adjustment ofa loading at the rotor assembly based on changing a vane pitch angle ofone or more vanes of a vane assembly of the aeronautical engine.

148. The computing system of any clause herein, wherein commanding theadjustment of the loading at the rotor assembly based on changing thevane pitch angle of one or more vanes of the vane assembly comprisescommanding a closing of the vanes to increase an amount of reversethrust generated by the aeronautical engine.

149. The computing system of any clause herein, wherein commanding theadjustment of the loading at the rotor assembly based on changing thevane pitch angle of one or more vanes of the vane assembly comprisescommanding an opening of the vanes to decrease an amount of reversethrust generated by the aeronautical engine.

150. A system for reducing noise generation for a single unducted rotorengine, the system comprising the propulsion system in accordance withone or more clauses of this chapter.

151. A propulsion system defining an engine centerline, the propulsionsystem comprising an unducted rotor assembly comprising a plurality ofblades extended radially relative to the engine centerline axis, therotor assembly configured to generate thrust substantiallyco-directional to the engine centerline axis, and a vane assemblypositioned aft of the rotor assembly, the vane assembly comprising aplurality of vanes extended radially relative to the engine centerlineaxis, wherein the propulsion system generates a power loading at therotor assembly of at least 25 horsepower per ft² at cruise altitude.

152. The propulsion system in accordance with one or more clauses ofthis chapter, wherein the propulsion system generates a power loading atthe rotor assembly of 100 horsepower per ft² or less at cruise altitude.

153. The propulsion system in accordance with one or more clauses ofthis chapter, wherein cruise altitude comprises an ambient air conditionbetween 4.85 psia and 2.14 psia.

154. A propulsion system of any one or more clauses herein configured toexecute the method of any one or more clauses herein.

What is claimed:
 1. A propulsion system defining an engine centerline,the propulsion system comprising: a rotor assembly configured to rotaterelative to the engine centerline axis, the rotor assembly comprising aplurality of blades, each blade of the plurality of blades configured torotate along a respective blade pitch angle axis; and a vane assemblypositioned in aerodynamic relationship with the rotor assembly, the vaneassembly comprising a plurality of vanes, each vane of the plurality ofvanes configured to rotate along a respective vane pitch angle axis; acontroller configured to execute operations, the operations comprising;moving each blade of the plurality of blades to a reverse thrustposition about its respective blade pitch axis, wherein a leading edgeof each blade is located aft of a trailing edge of the respective bladeat a radial span location when in the reverse thrust position; andadjusting each vane of the plurality of vanes about its respective vanepitch axis when the plurality of blades is in the reverse thrustposition to modify an amount of reverse thrust generated by thepropulsion system.
 2. The propulsion system of claim 1, wherein therotor assembly is unducted.
 3. The propulsion system of claim 1, whereinthe one or more blades is configured to generate forward flow over afirst portion of a blade span, and wherein the one or more blades isconfigured to generate reverse flow over a second portion of the bladespan.
 4. The propulsion system of claim 1, wherein the one or moreblades is configured to generate forward flow below 50% of a blade span,and wherein the one or more blades is configured to generate reverseflow at or above 50% of the blade span.
 5. The propulsion system ofclaim 1, wherein adjusting each vane of the plurality of vanes about itsrespective vane pitch axis when the plurality of blades are in thereverse thrust position comprises rotating one or more vanes along thevane pitch axis up to 15 degrees open or up to 15 degrees closed from adesign point.
 6. The propulsion system of claim 1, wherein adjustingeach vane of the plurality of vanes about its respective vane pitch axiswhen the plurality of blades are in the reverse thrust positioncomprises rotating one or more vanes along the vane pitch axis up to 10degrees open or up to 10 degrees closed from a design point.
 7. Thepropulsion system of claim 1, wherein adjusting each vane of theplurality of vanes about its respective vane pitch axis when theplurality of blades are in the reverse thrust position comprisesrotating one or more vanes along the vane pitch axis up to 5 degreesopen or up to 5 degrees closed from a design point.
 8. The propulsionsystem of claim 1, wherein adjusting each vane of the plurality of vanesabout its respective vane pitch axis when the plurality of blades is inthe reverse thrust position comprises closing the vanes to increase theamount of reverse thrust generated by the propulsion system.
 9. Thepropulsion system of claim 1, wherein adjusting each vane of theplurality of vanes about its respective vane pitch axis when theplurality of blades is in the reverse thrust position comprises openingthe vanes to decrease the amount of reverse thrust generated by thepropulsion system.
 10. The propulsion system of claim 1, wherein therotor assembly is ducted.
 11. A method for generating reverse thrust fora single stage unducted rotor engine with a vane assembly positioned inaerodynamic relationship, the method comprising: adjusting a blade pitchangle at one or more blades of the rotor assembly to position a bladeleading edge aft of a blade trailing edge at a radial span location; andadjusting loading at the rotor assembly based on changing a vane pitchangle of one or more vanes of the vane assembly.
 12. The method of claim11, wherein adjusting loading at the rotor assembly based on changingthe vane pitch angle of one or more vanes of the vane assembly comprisesclosing the vanes to increase the amount of reverse thrust generated bythe engine.
 13. The method of claim 11, wherein adjusting loading at therotor assembly based on changing the vane pitch angle of one or morevanes of the vane assembly comprises opening the vanes to decrease theamount of reverse thrust generated by the propulsion system.
 14. Themethod of claim 11, wherein adjusting loading at the rotor assemblybased on changing the vane pitch angle of one or more vanes of the vaneassembly comprises rotating one or more vanes along the vane pitch axisup to 15 degrees open or up to 15 degrees closed from a design point.15. The method of claim 11, wherein adjusting loading at the rotorassembly based on changing the vane pitch angle of one or more vanes ofthe vane assembly comprises rotating one or more vanes along the vanepitch axis up to 10 degrees open or up to 10 degrees closed from adesign point.
 16. The method of claim 11, wherein adjusting loading atthe rotor assembly based on changing the vane pitch angle of one or morevanes of the vane assembly comprises rotating one or more vanes alongthe vane pitch axis up to 5 degrees open or up to 5 degrees closed froma design point.
 17. The method of claim 11, wherein adjusting the bladepitch angle at one or more blades of the rotor assembly to position theblade leading edge aft of the blade trailing edge at the radial spanlocation comprises: generating forward flow below 50% of a blade span;and generating reverse flow at or above 50% of the blade span.
 18. Acomputing system, the computing system configured to store instructionsthat, when executed by the one or more processors, performs operations,the operations comprising: commanding an adjustment of a blade pitchangle at one or more blades of a rotor assembly of an aeronauticalengine to position a blade leading edge aft of a blade trailing edge ata radial span location; and commanding an adjustment of a loading at therotor assembly based on changing a vane pitch angle of one or more vanesof a vane assembly of the aeronautical engine.
 19. The computing systemof claim 18, wherein commanding the adjustment of the loading at therotor assembly based on changing the vane pitch angle of one or morevanes of the vane assembly comprises commanding a closing of the vanesto increase an amount of reverse thrust generated by the aeronauticalengine.
 20. The computing system of claim 18, wherein commanding theadjustment of the loading at the rotor assembly based on changing thevane pitch angle of one or more vanes of the vane assembly comprisescommanding an opening of the vanes to decrease an amount of reversethrust generated by the aeronautical engine.